Poster Session Wednesday, June 15 from 5:00 to 7:00 PMippw2016.jhuapl.edu/docs/abstracts/Complete_Poster_session_abstr… · Poster Session Wednesday, June 15 from 5:00 to 7:00 PM

Poster Session

Wednesday, June 15 from 5:00 to 7:00 PM

PLANETARY OBJECT GEOPHYSICAL OBSERVER

(POGO) FOR IN SITU OBSERVATIONS OF AIRLESS

BODIES

Airless Bodies Elena Adams [email protected]

MOBILITY TECHNOLOGY FOR ENCELADUS

TERRAIN ACCESSIBILITY (ETNA)

Airless Bodies Paul Witsberger [email protected]

MOBILE UTILITY PLATFORM FOR PROBING

EUROPA'S TERRAIN (MUPPET)

Airless Bodies Benjamin Libben [email protected]

Mission on Autonomous UAV for Lunar

Inspection

Airless Bodies BALAJI SOUNDARARAJAN [email protected]

Optimal Electrode Arrangements for Electrostatic

Alleviation of Reentry Blackout

Cross Cutting Technologies Siddharth

Krishnamoorthy, Sigrid

Close

skrishnamoorthy@stanfor

d.edu

Magnetohydrodynamically Enhanced

Deceleration for Planetary Entry Vehicles

Current and Proposed EDL

Technology

Hisham K. Ali [email protected]

A Decelerator for Pluto Entry, Descent, and

Landing


Technology

Benjamin Goldman benjamin.d.goldman@gae

rospace.com

ASSESSMENT OF SYSTEM-LEVEL MASS AND

PERFORMANCE OF ENTRY, DESCENT, AND

LANDING TECHNOLOGIES FOR HUMAN AND

ROBOTICS MISSIONS TO MARS


Technology

Eiji Shibata, Ben Libben,

Maxim de Jong, Sarag

Saikia

[email protected]

Combining Camera-based Hazard Detection and

Terrain Relative Navigation in a SLAM-like

approach


Technology

Svenja Woicke and Erwin

Mooij

[email protected]

Inflatable Decelerators for Human-Scale Missions

to Mars


Technology

Robert Dillman [email protected]

v

Lifting Entry Validation Experiment (LEVE) for

Lifting Entry Atmospheric Flight (LEAF) Systems


Technology

Greg Lee [email protected]

Magnetohydrodynamic Energy Generation and

Atmospheric Breathing Supersonic

Retropropulsion for Mars Descent


Technology

Keir Gonyea [email protected]

Mars 2020 Terrain Relative Navigation

Accommodation


Technology

Aaron Stehura [email protected]

SENSOR DATA FUSION FOR HAZARD MAPPING

AND PILOTING


Technology

Kanani [email protected]

m

Technology Overview and Assessment for Small-

Scale EDL Systems


Technology

Casey Heidrich [email protected]

Development of Atmospheric Entry Trajectory

and Aerothermodynamics Code & Design of

Deployable Aer-odynamic Decelerator for

Sample Return Mission

Drag, Aerobraking and

Aerocapture Technologies

Arturs Jasjukevics [email protected]

Gas Generators and their potential to support

Human-Scale HIADs



Richard Bodkin [email protected]

The Horizon 2020 project IRENA (International

Re-Entry demoNstrator Action)



Jean-Marc Bouilly jean-

[email protected]

s.net

Air Dust Removal Tool (AirDRT): A Novel Concept

for Cleaning Rocks and Spacecraft Surfaces off

Dust on Mars Using High Speed Propeller

Instrumentation and Science

Investigations

Kris Zacny zacny@honeybeerobotics.

com

Astrobiology at CNES: research and development

programmes paving the way for the future.


Investigations

Pascale Chazalnoel [email protected]

r

CHARACTERIZATION OF A METHOD FOR INVERSE

HEAT TRANSFER


Investigations

M. E. Pizzo, K. S. Bey, and

D. E. Glass

[email protected]

MARSDROP: GETTING MINIATURE INSTRUMENTS

TO THE SURFACE OF MARS AS SECONDARY

PAYLOADS


Investigations

Robert Staehle [email protected].

gov

IPPW-13 Program Abstracts - Poster Session

July 13-17, 2016 Laurel MD USA http://ippw2016.jhuapl.edu/ 1

ULTRASONICALLY ASSISTED PENETRATION

THROUGH GRANULAR MATERIALS


Investigations

David Firstbrook [email protected]

a.ac.uk

INVESTIGATING TITAN’S GEOLOGY AS MEANS

FOR THE PREPARATION OF FUTURE MISSIONS

Landing Site Selection and

Definition

Anezina Solomonidou Anezina.Solomonidou@jpl.

nasa.gov

A MULTI-PLANET, MULTI-SPACECRAFT FLAGSHIP

CLASS MISSION CONCEPT TO EXPLORE A GAS

GIANT AND AN ICE GIANT PLANET

Missions A. Mudek and S. J. Saikia [email protected]

A Summary Of The SOAREX-8 and SOAREX-9 Sub-

orbital Test Flights

Missions Marcus S. Murbach [email protected]

ov

Assessment of Venus Aerial Platforms for In-Situ

Measurements and Analysis of Optimal Altitude

for Superpressure Balloons

Missions Robert Rolley [email protected]

ATHENEA - A mission proposal to a Main Belt

Comet

Missions C. Mockel [email protected]

Biosignature Explorer for Europa (BEE) Probe –

Directly Searching for Life Evidence on Europa

Missions Michael Amato [email protected]

Concept of Lunar Ballistic Robot Exploration

(BARE)

Missions BALAJI SOUNDARARAJAN [email protected]

Cubesat Application for Planetary Entry (CAPE)

Missions: Micro-Return Capsule (MIRCA)

Missions Jaime Esper [email protected]

Exomars Heatshield Missions Yann Mignot [email protected]

s.net

GLIDING THROUGH THE HYDROCARBON LAKES

OF TITAN USING A STEERABLE PARACHUTE.

Missions Ye Lu [email protected]

Gravity Tractor Dynamics under the Effect of

Non-Homogenous Gravity Fields of Asteroids

Missions S. P. Shekhar and Y.

Ketema

[email protected]

INTRODUCTION AND PROGRESS OF THE AIM

CUBESAT OPPORTUNITIES, (COPINS).

Missions D. Binns [email protected]

IXV: THE MISSION OF THE FIRST LIFTING BODY

REENTRY VEHICLE

Missions Davide Bonetti davide.bonetti@deimos-

space.com

MARIUS MISSION: Proposal for an ESA M-class

seismology mission to Europa.

Missions Mathijs Van de Poel mathijs.vandepoel@gmail.

com

KRUPS Design Integration Overview Missions Chris Meek [email protected]

RAPID EXPLORATION OF THE SOLAR SYSTEM

USING AEROGRAVITY-ASSIST

Missions Peter Edelman, Eiji

Shibata, Sarag Saikia, Jim

Longuski

[email protected]

ROBOTIC GRIPPER DESIGN AND TESTING FOR

POTENTIAL MARS SAMPLE RETURN (MSR)

Missions RYAN MCCORMICK [email protected]

Saturn-Uranus Trajectories for Multi-Planet

Missions

Missions Kyle Hughes [email protected]

PLANETARY PROBE ENTRY ATMOSPHERE

RECONSTRUCTION USING SYNTHETIC AIR DATA

SYSTEM

Modeling, Simulation & Testing Chris Karlgaard [email protected]

GEERGEER: A unique capability to support

planetary exploration and probe development

Modeling, Simulation & Testing Tibor Kremic [email protected]

Hypervelocity Expansion Tube Studies of Blunt

Body Aerothermodynamics in CO2.

Modeling, Simulation & Testing Matthew Leibowitz [email protected]

LOW DENSITY SUPERSONIC DECELERATOR (LDSD)

SUPERSONIC FLIGHT DYNAMICS TEST (SFDT)

PLUME INDUCED ENVIRONMENT MODELLING

Modeling, Simulation & Testing Brandon L. Mobley [email protected]

ov

SIZING METHODS FOR ADVANCED MARS ENTRY

DESCENT AND LANDING SYSTEMS

Modeling, Simulation & Testing Marcus Lobbia [email protected].

gov

SurRender, an image rendering software for

scientific space scene simulation

Modeling, Simulation & Testing Roland Brochard [email protected]

om



PLANETARY OBJECT GEOPHYSICAL OBSERVER (POGO) FOR IN SITU OBSERVATIONS OF

AIRLESS BODIES. E. Adams1, E. Hohlfeld

1, S. Murchie

1, S. Liang

1, B. Williams

1, A. Sharma

1, J. Kelley

1, D.

Ponnusamy1, R. Hourani

1, E. Donald

1, R. Hearty

1, J. Neville

1, K. Lake

1, D. Myers

1, S. Gaumier

1, C. Lopez

1, J.

Ramirez1, J. Daniels

1, S. Cheng

1, R. Hacala

1, B. Stilwell

1, R. Lorenz

1, P. Peplowski

1, J. Goldsten

1 1The Johns Hop-

kins Applied Physics Laboratory (11000 Johns Hopkins Road, Laurel MD, 20723-6099, [email protected])

Planetary Object Geophysical Observer (POGO) is

a hopping lander that can be used on small planetary

bodies (e.g. comets, asteroids, and small moons) [1, 2].

It relocates on the surface by actuating a simple voice

coil mechanism. The current payload samples the sur-

face elemental composition using a gamma-ray spec-

trometer and an alpha particle x-ray spectrometer, and

the regolith properties using accelerometers. POGO is

unique because it can survive landing velocities up to 5

m/s allowing for deployment at high altitude, can ac-

quire science measurements in any landed orientation

[3], and its hopping distances vary based on current

input to the mechanism (e.g. on Phobos, it can hop

~10s of meters; longer distances at smaller bodies). It

operates on a pre-canned sequence, performing a ran-

dom walk on the surface in between periods of meas-

urements. POGO is constantly transmitting the data on

a loop to the orbiting spacecraft, allowing for simple

operations. Using primary batteries as its power source,

it can land at any latitude, sun or shade, and for the

current payload, operate on the surface ~5 days. POGO

full scale prototype was developed and tested to TRL 5

by JHU/APL as part of internal research and develop-

ment, and has since been funded by NASA ARM grant

and by NASA New Frontiers Homesteader grant. We

will describe the POGO technology development to

date, and current qualification path to TRL 6.

References:

[1] Hill, S. et al., (2014), International Workshop

on Instrumentation, GSFC, #1039. [2] Adams, E. et al.

(2016) IEEE Aerospace Conference, Big Sky, MO. [3]

Nunez et al. (2014) , International Workshop in In-

strumentation, GSFC, #1145.



MOBILITY TECHNOLOGY FOR ENCELADUS TERRAIN ACCESSIBILITY (ETNA) P. A. Witsberger1†,

L. P. Podesta2†, E. Shibata3†, R. Potter4†, A. J. Anderson5‡, K. W. C. Coleman6†, S. J. Fulton5†, C. Guven5†, S. C. Ho6†,

N. Hobar5†, E. E. Hock5†, L. Klug6†, R. Lucas5†, P. Machuca6†, S. Matlock5†, A. Mudek6†, J. L. Pouplin6†, C. M.

Reinert6†, R. J. Rolley5‡, D. Sharpless5†, and S. J. Saikia7†, 1Undergraduate Student, [email protected], 2Graduate

Student, [email protected], 3Graduate Student, [email protected], 4Graduate Student, [email protected], 5Undergraduate Student, 6Graduate Student, 7Visiting Assistant Professor, [email protected], †School of Aero-

nautics and Astronautics, Purdue University, 701 W. Stadium Ave., IN, 47907, ‡School of Mechanical Engineering,

Purdue University, 585 Purdue Mall, West Lafayette, IN, 47907.

Introduction: Far from our warm world, there exist

icy moons orbiting Jupiter and Saturn. Of all the icy

moons, the Planetary Decadal Survey identifies two (Ju-

piter’s moon Europa, and Saturn’s moon Enceladus) for

their potential to harbor life [1]. On their surfaces, Eu-

ropa and Enceladus are dominated by icy fea-

tures. However, there are many indicators that point to

subsurface oceans beneath their frozen exteriors, which

present all the fundamental necessities for the formation

of life. The most recent Planetary Decadal Survey iden-

tifies several key features that are of interest on these icy

moons [1]. Therefore, the next step in the exploration of

Enceladus is a mobile lander which can traverse the

rough terrain and make in situ measurements at multiple

locations of high-scientific interest. The exploration of

Enceladus presents many challenges similar to those on

the other icy moons such as regions varying from slick

and smooth to quite rugged with high cliffs, deep cre-

vasses, blocky ice boulders, and even tall penitentes.

Current state of the art technology involves using rov-

ers, such as the MSL rover [2]. However, the rover de-

sign has a maximum terrain height, which prevents it

from being able to navigate large portions of Enceladus.

The 2013–2022 Vision and Voyages for Planetary

Science identifies three key questions for any planetary

satellite [1]:

How did the satellites of the outer solar system

form and evolve?

What processes control the present-day behavior of

these bodies?

What are the processes that result in habitable envi-

ronments?

For all three of those questions, using purely ground-

based science observations is not enough. In situ meas-

urements must be made with some form of mobility to

increase the science potential. For an Enceladus mis -

sion, the following science goals are listed in the Deca-

dal Survey in order of priority [1]:

What is the nature of Enceladus’s cryovolcanic ac-

tivity?

What are the internal structure and chemistry of En-

celadus?

What is the nature of Enceladus ’s geologic history?

How does Enceladus interact with the rest of the

Saturnian system?

What is the nature of the surfaces and interiors of

the other moons?

Why Enceladus?: The Cassini space probe col-

lected data that suggests the potential existence of com-

plex organic molecules, heat, and liquid oceans beneath

the South Pole of Enceladus. The evidence of organic

molecules was collected during a fly-through of plumes

of water ejected from cryovolcanoes. Anomalous mag-

netometer data collected during these flybys suggested

there was a pocket of fluid beneath the ice [1]. The co-

existence of cryovolcanic activity, the presence of or-

ganic molecules, and the possibility of liquid water sug-

gest Enceladus may support life.

Fig. 1: Plume mass spectrum from the Cassini ion and neu-

tral mass spectrometer [1].

The biological potential of Enceladus is the most in-

triguing area of scientific study. Specifically, future

missions should seek to understand whether the condi-

tions for life are present on Enceladus. Science objec-

tives include identifying the location and distribution of

liquid water, characterizing the physical conditions of

the active Tiger Stripe regions, analyzing the chemical

and biogenic structures present in the plumes, and char-

acterizing the surface chemistry and global distribution

of chemical species. A lander on Enceladus would be

able to conduct these chemical measurements at multi-

ple locations on the surface, sample the plumes to deter-

mine their constituents, and characterize the motion of

tectonic features and the location and distribution of liq-

uid water [3].



How Can We Do It?: The region of primary scien-

tific interest, the South Pole region of Enceladus, has a

very rugged terrain. Ridges and cliff faces up to 500 m

tall [5], large ice blocks up to 100 m in diameter [6], and

several massive rifts called sulci (the Tiger Stripes) that

are on the order of 100 km long, 2 km wide, and 500 m

deep from which cryovolcanic plumes originate [7]

characterize the South Pole region. Because of the rough

exterior, a unique mobility system is required in order

to study distinct areas and geographic formations on the

surface. The surface of Enceladus near the plumes is

coated in a layer of fine water-ice particles that originate

from the plumes and is of high scientific value [5].

Fig. 2: Geyser locations along “Tiger Stripe” fractures [4].

We propose a conceptual design of a mobility sys-

tem to search for evidence of life on Enceladus. The

main phases of this study are the exploration of the trade

spaces involved in novel mobility systems, develop-

ment of these concepts, and finally comparing and eval-

uating an optimal design. The final output of this study

is a mission architecture from landing to end of life and

would focus on the conceptual subsystem design of the

following: mobility system, power supply and manage-

ment, science instrument suite selection, telecommuni-

cations, and thermal control system.

Initially, the Science Traceability Matrix is drawn

up in order to identify possible landing sites and relevant

instruments. We later collect information on heritage

and novel mobility systems, assess their merits and

shortcomings, and consider possible combinations that

could complement each other. Once a broad view of

various systems is gathered, trade studies are performed

to evaluate the versatility and science return of each mo-

bility system in order to objectively compare them and

deduce the optimal configuration.

Point design is the final step of this process, which

entails selecting a sensor suite, developing the selected

mobility system in depth for the chosen mission profile,

and specifying all necessary subsystems with their com-

puted requirements. To ensure successful integration,

concepts of each specific subsystem are developed in

detail. Conceptual subsystem design also involves trade

studies specifically for spacecraft resources (space

availability, power requirements, structural considera-

tions).

Fig. 3: Conceptual rendering of terrain around geysers on the

surface of Enceladus [8].

References: [1] Squyres S. W. (2011) “The NRC

Planetary Decadal Survey” 2011 Aerospace Confer-

ence. [2] Heverly M. et al. (2013) Journal of Field Ro-

botics, Vol. 30, 6, 835-846. [3] Bly V. et al (2007) “En-

celadus Flagship Mission Concept Study”. [4] Porco

C. C. et al (2014) The Astronomical Journal, 148(3), 45.

[5] Spencer J. R. (2009 Saturn from Cassini-Huygens,

Springer, pp. 683-724. [6] Martens H. R. (2015) Icarus,

245, pp. 162-176. [7] Porco C. C. (2006) Science, 311,

5766, pp. 1393-1401. [8] Carroll, M. W., and Lopes, R.

M. C., Alien seas: oceans in space.



MOBILE UTILITY PLATFORM FOR PROBING EUROPA'S TERRAIN (MUPPET) B. Libben1†, B. Tack-ett2†, N. Kolencherry3†, A. Sharma4†, Y. Lu5†, A. Bruce5†, J. Mansell5†, G. Smith5†, T. Ukai5†, M. Vanka5†, S. Vutu-kuri5†, A. Zhang5‡, A. Arora6†, K. Bush6†, J. Lepak6†, J. Millane6†, B. Miller6†, L. Moroz6†, T. Murray6†, B. Smith6†, M. Young6†, and S. Saikia7†, 1Graduate Student email: [email protected], 2Graduate Student email: [email protected], 3Graduate Student email: [email protected], 4Graduate Student email: [email protected], 5Grad-uate Student, 6Undergraduate Student, 7Visiting Assistant Professor, †School of Aeronautics and Astronautics, Purdue University, 701 W. Stadium Ave., West Lafayette, IN, 47907-2045, ‡School of Interdisciplinary Engineering, Purdue University, 701 W. Stadium Ave, West Lafayette, IN, 47907-2045

Introduction: NASA’s Planetary Decadal Survey

highlights the icy moons of Jupiter and Saturn as the most logical places in the solar system to search for life. Jupiter’s icy moon, Europa, has been of particular inter-est lately in study of potentially habitable environments [1]. These environments are quite extreme, and could potentially harbor microscopic lifeforms known as ex-tremophiles which can thrive in such habitats [2]. The temperature on the surface of Europa is extremely cold, which is why the theorized sub-surface ocean is of greater interest in the search for life, than the surface. However, before committing vast resources based on limited information to exploring this possible sub-sur-face ocean, which can be up to tens of kilometers under the icy crust, the surface of this icy moon requires fur-ther exploration. Surface exploration can provide a deeper understanding of the moon's composition and provide further insight for future sub-surface ocean ex-ploration activities, helping lower both cost and risk.

Figure 1: Artist's depiction of calving on Europa. Macula features in choas regions are thought to be caused by calving, similar to events seen on Earth. Choas regions with evidence of recent resur-facing can provide data on sub-surface ice and water features [7].

The objective of this paper is to design a mobility system with the purpose of exploring the surface of Eu-ropa for particular science objectives including the search for life.

The Vision and Voyagers for Planetary Science De-cadal Survey for 2013-2022 [1] posed the following questions regarding studies of planetary satellites: How did the satellites of the outer solar system

form and evolve? What processes control the present-day behav-ior of these bodies?

What are the processes that result in habitable environments?

The exploration of Europa is expected to help an-swer these questions about the formation of the icy moons, the interactions that result in their current be-havior, and the search for habitable environments be-yond Earth.

Proposed Missions: NASA JPL’s concept Europa Multiple-Flyby mission includes an orbiter spacecraft and a possible lander to investigate whether the icy moon could harbor conditions suitable for life [5]. The mission will provide greater resolution of surface fea-tures for future exploration. NASA also did a Europa lander study that featured a soft lander on the surface to perform in situ analysis [5]. The lander will carry six instruments for it investigations, and will be incapable of mobility.

While static landers involve less risk in the overall mission, the benefits of an active mobility system on Europa’s surface can outweight the increased complex-ity. A mobility system is capable of landing in a safe location and moving to an area of scientific interest that may be too hazardous to land in.

Science, Science, Science: The overall goals of a first Europa mission are to assess habitability, detect life, and gather planetary science data [1]. The scientific objectives are chosen to be measuring the magnetic field, radiation, surface composition, and plate tecton-ics, as well as verifying the existence of a sub-surface ocean and detection of life. Most of these objectives can be satisfied using already developed technologies from missions such as MSL, Philae, and ExoMars. Some of these instruments have already been designed to func-tion in similar environments to those on Europa’s sur-face. The final selection of instruments are designed or selected, based on mass, power, volume, environment, and technology constraints.

Setting up Basecamp: The landing site selection criteria for such a system is paramount and is chosen to provide access to the most scientifically important sites while minimizing the impact of high radiation and ex-treme surface features which is characteristic of Europa [3]. Europa has quite a few points of interest as it is cur-rently unexplored, except for fly-by imagery from Gal-ileo [4].



Figure 2: Highest available resolution image of Europa's surface taken by Galileo (6m/pixel) while 600 km above the surface. The pic-ture is not re-projected, so the view is the same as if looking out a plane window. The image shows rough terrain throughout, with many of the features such as crevices, steep slopes, and other chaotic fea-tures [3].

The surface features of Europa have been catego-rized as linea, flexus, macula, and chaos regions based on the images taken by the Galileo spacecraft. For an initial exploration mission to Europa, selecting a land-ing site in an area close to one of these surface features can increase the science potential of the mission [5]. To mimize radiation exposure, a landing site in the leading hemisphere of the moon will be selected [3].

Figure 2: Blue contours show radiation intensity on Europa’s surface, labeled by the energy intensity, and how deeply they penetrate the sur-face. (excluding the effects of secondary particles) [6].

The southern region of Europa is also of interest as it is hypothesized to be the most likely location of hy-drothermal vents [3]. Surface challenges near features of interest include craters, large boulders, crevices, mountains, penitentes, and steep inclines/declines.

Phoning Home: Landing site selection also has a direct impact on the choice of the communication pack-age that the mobility system will carry. Communication strategies using either direct communication to Earth or a Jupiter orbiter as a relay are explored. Large surface features such as mountains or close boulders will limit the time window the mobility system can send and re-ceive data packets. Eclipses caused by Jupiter and the Sun are also factors when designing the communication link.

Current Mobility Systems: The surface challenges near features of interest provide significant obstacles for

state of art mobility systems. The wheel and axle sys-tems used in traditional rovers will not be very effective at traversing icy terrain with multiple cracks, boulders, and penitentes. Tread based designs such as snowmo-biles are effective at traversing loose terrain, but they struggle to overcome the same obstacles inhibiting a wheel and axle system.

Going Mobile: To meet the science objectives, the mobility system is designed to take measurements at four different points on the surface that are at least one kilometer apart. Mobility platforms for performing both spatial and temporal coverage of Europa's high interest surface features can be classified into two broad classes – Active and Static systems. An active mobility system achieves motion via thrust or motor devices with the ability to move from one location to another. A static mobility system is a dispersion of probes with a central lander that does not move. Hybrid systems can be ob-tained by combining aspects from both mobility types.

Selection Process: Science requirements and the landing site are selected to meet NASA’s goals for ex-ploring Europa. Formal concept generation and selec-tion methods are applied to the preliminary designs to down-select and refine the mission architecture. These final designs are then sized to meet the mission require-ments, and based on their mass, power requirements, and overall volume, an optimal solution is selected for the mission.

References: [1] Board, Space Studies. (2012) Vision and Voy-

ages for Planetary Science in the Decade 2013-2022. [2] Thomas, D. N., and G. S. Dieckmann. (2002) Ant-arctic sea ice--a habitat for extremophiles." Science 295, no. 5555, 641-644. [3] Pappalardo, R. T., et al. (2013) Science potential from a Europa lander. Astro-biology 13.8, 740-773. [4] Europa Photo Gallery. NSSDCA. [5] JPL (2012) Europa Study 2012 Report. D-71990 [6] Patterson, G. W., et al. (2012) Character-izing Electron Bombardment of Europa’s Surface by Location and Depth. Icarus, 220:286–290. [7] Schmidt, B. Europa’s Great Lakes (Image). [8]



Mission on Autonomous UAV for Lunar Inspection.

Balaji Soundararajan

Technical University of Berlin, Marchstr 12-14, 10587, Berlin, Germany, email: [email protected]

Introduction: The motivation of this paper is to

present about the possible expansion of the lunar ex-

ploration capabilities with or without Human presence

on the Moon using an Unmanned Aerial Vehicle

(UAV). With major Space organisations around the

world wanting to go back to the Moon in 2020s, In-situ

Resource Utilization (ISRU) will be an important topic

for future. Exploration of diverse locations on the sur-

face of the moon will help us in understanding the

moon better, as we haven’t gone to places like the far

side, inside deep craters and Paleo Regoliths (which

are preserved under lava beds). The exploration of

these hard areas can be highly possible using Liquid

Hydrogen (LH2) powered UAV. As the system limita-

tions of surface based Lunar rovers limit them to not

travel beyond 40Km, usage of UAVs will be highly

helpful. The UAV, which has thrust vectoring capabil-

ity, shall be designed to go as far as 500 Km round trip

from the landing site or from the lunar outpost. Initial-

ly, the fuel is carried along with the crew or with the

landing module. As it is possible to extract hydrogen

from the lunar regolith, we can generate LH2 for later

use. Because of the fact that the moon has very low

gravity (1.622 m/s²), meagre amount of fuel will help

us achieving a sortie.

Basic design. Considering the UAV in a cylindri-

cal shape (1m long and 0.6 m in diameter), Basic cal-

culations reveal that with 25Kg of fuel, it is possible to

travel 500 Km. The design can be customized as per

requirements. The UAV is divided into two halves,

with the top compartment housing the avionics and

fuel and bottom compartment having provision for

sample collection or other payload.

Figure.1 Lunar UAV with Door openened for

sample collection

The bottom area has a unique door with blade tip.

Once the UAV lands or hovers at a specified location,

it collects the samples. Pulsed thrust is generated by

the thruster to generate vibrations in order to break

open any hard regolith layers. The Thrust vector noz-

zles are deviated upwards during sample collection so

as to avoid contamination of samples that are being

collected. The UAV’s speed is limited to 2.00 km/sec

as it cannot be more than Moon’s escape velocity of

2.38 Km/sec. due to the absence of atmosphere, it is

possible to achieve these speeds. The UAV has a HD

video camera at the front and relays high definition

video to the L2 Lagrange point or to the Lunar outpost.

Avionics. The UAV has an advanced avionics ca-

pability such as computer vision and image recognition

for feature detection. This helps in understanding the

lunar environment and acts precisely. A LIDAR based

technique for for high resolution maps is used to assess

the sites. Guidance, navigation and autonomous con-

trol with provision to change to manual and semi au-

tomatic mode can be done. A RSSI based tracking en-

vironment for outdoor environments is also used. A

collision avoidance system will make sure UAV

doesn’t hit any where. A radio communication using

software defined radio is used to send vital data back

to Lunar outpost or station at L2.

Possible applications. The Lunar UAV is highly

handy during exploration. It can be used as early warn-

ing system during lunar colonization. It also assess the

possible exploration sites, collects samples such as

regolith, lava layers and ice crystals. Launching of

small, cubesats into the lunar orbit is possible from this

UAV.

Conclusion. With advancement in technology and

the advent of Earth based UAVs, Lunar UAV will

soon be a possible thing. Due to their system design,

ability to fly across swiftly covering more distances,

effortlessly going to hard areas where rovers cannot be

sent easily, makes it more interesting than convention-

al rovers. This will surly prove to be an Exploration

friendly vehicle.



OPTIMAL ELECTRODE ARRANGEMENTS FOR ELECTROSTATIC ALLEVIATION

OF REENTRY BLACKOUT. S. Krishnamoorthy1 and S. Close2

1PhD. Candidate, Department of Aeronautics and Astronautics, Stanford University, Stanford, CA,

USA. Email address: [email protected] 2Assistant Professor, Department of Aeronautics and Astronautics, Stanford University, Stanford,

CA, USA. Email address: [email protected]

Introduction: The reentry blackout

phenomenon affects most spacecraft reentering a

planetary atmosphere from the vacuum of space.

The blackout is caused by the presence of highly

mobile electrons in a plasma layer that surrounds

the vehicle. The electrons in the plasma layer

prevent the transmission of electromagnetic

waves below the plasma frequency, leading to a

loss of command, communication and telemetry

signals during the blackout period. The blackout

period may last up to several minutes, and is a

major contributor to the landing ellipse. In the

context of human spaceflight, this may also

present a significant safety hazard. Recently,

Krishnamoorthy et al. [1] [2] suggested a possible

method to alleviate reentry blackout by using

strong pulses of negative voltage, called Pulsed

Electrostatic Manipulation (PEM). Their

analysis, resulting from a 2-D electrostatic

Particle-in-Cell (PIC) simulation, showed the

clearance of electrons up to three electrode

lengths perpendicular to the vehicle surface. This

analysis was conducted using a two dimensional

simulation. It is a well-known fact that for a given

voltage, the electric field produced strongly

depends on the shape of the equipotential

surfaces, which in turn depend on the shape of the

electrode [3]. Thus, it is reasonable to expect that

the performance of the blackout alleviation

system is strongly dependent on the shape of the

electrodes used to apply the voltage and also the

placement of the electrodes on the spacecraft

surface. The aim of the current work is to

investigate the dependence of system

performance on these two parameters.

Pulsed Electrostatic Manipulation

(PEM): When a plasma is subject to an

electronegative pulse, the electrons get repelled

from the vicinity of the electrode, while the ions

accumulate near it, in order to shield out the

voltage applied. The mobility of the electrons is

much higher than that of the ions, which creates

an area devoid of electrons till the ions move in

to shield the applied voltage [4].

(a)

(b)

(c)

Figure 1: Schematic descripting of PEM for reentry

blackout alleviation. (a) Locations on spacecraft (b)

Side view of one of the electrodes (c) Top view of one

possible electrode arrangement



mailto:[email protected]


This period of time over which the electron

clearance occurs creates a “window” for

communication. Eventually, the ions shield out

the potential, and eventually cause dielectric

breakdown; this mandates that the pulse be turned

off. The technique is schematically described in

Fig. 1. The electrodes are placed at antenna

locations, insulated from the spacecraft surface

and the plasma. Electronegative pulses are

applied through these electrodes, locally reducing

plasma density, allowing for radio

communication from the vehicle.

Simulation Setup: The electron content in

the vicinity of the electrodes is severely depleted,

leading to very low electron densities. Further,

the process of electron clearance is highly

dynamic, requiring the use of time-accurate

simulations. Therefore, the interaction of the

plasma with the electrodes is simulated using the

PIC algorithm, described in detail in [5]. A two

dimensional analysis of this problem was

presented in [1] and [2]. While the 2-D analysis

enables a good preliminary study of system

performance, it is impossible to study the effect

of electrode shape and placement; a full 3-D

simulation is needed for this task. However, the

extension of the simulation from 2-D to 3-D

poses several challenges to numerical efficiency,

which must be tackled in order to perform the

simulations. These challenges are tackled with

multi-grid methods for the solution of Poisson

equations and parallel computing to process

particle dynamics. The PICard Particle-in-Cell

solver (written and tested at Stanford University)

is equipped with these computational techniques

and was used for the results to be presented in this

paper. A two dimensional cross section of the full

domain used for 3-D analysis is shown in Fig 2.

Figure 2: A 2-D cross section of the full 3-D

simulation domain. Plasma-dielectric interaction is

modeled through special boundary conditions [2],

farfield boundaries are modeled as zero-field open

boundaries.

Projected Results: Using the PICard

simulation software, the time-accurate behavior

of the communication window will be simulated

for different electrode shapes. The results from

these simulations will be used to study the

variation of system performance parameters

(such as communication window size and

dielectric breakdown risk) with the

characteristics of the applied voltage pulse.

Further, the effect of electrode placement on

system performance will also be discussed.

References

[1] S. Krishnamoorthy and S. Close, "Particle in

Cell (PIC) simulations of plasma-electrode

interactions for Reentry Blackout

Alleviation," in 46th AIAA Plasmadynamics

and Lasers Conference, Dallas, TX, 2015.

[2] S. Krishnamoorthy and S. Close,

"Investigation of plasma-surface interaction

effects on Pulsed Electrostatic

Manipulation for reentry blackout

alleviation," (Under Review) Journal of

Physics D: Applied Physics, 2016.

[3] D. J. Griffiths, Introduction to

Electrodynamics (4th Ed.), Pearson

Education, 2014.



[4] S. G. Bilen and B. E. Gilchrist, "Transient

plasma sheath model for thin conductors

excited by negative high voltage with

application to electrodynamic tethers,"

IEEE Transactions on Plasma Science, vol.

28, no. 6, 2000.

[5] C. K. Birdsall and A. B. Langdon, Plasma

Physics via Computer Simulation, CRC

Press, 2004.



MAGNETOHYDRODYNAMICALLY ENHANCED DECELERATION FOR PLANETARY ENTRY

VEHICLES. H. K. Ali1* and R. D. Braun2*, 1Graduate Research Assistant, [email protected], 2David and An-

drew Lewis Professor of Space Technology, [email protected], *Space Systems Design Lab, Daniel Gug-

genheim School of Aerospace Engineering, Georgia Institute of Technology, 270 Ferst Drive, Atlanta, GA 30331.

Introduction: Future missions to Mars such as a

Mars sample return mission and potential human mis-

sion will require much higher masses than have ever

been landed on Mars. Previous Mars missions have re-

lied primarily on Viking era technology for entry de-

scent and landing [1].The limit of this technology is be-

ing reached, with the Mars Science Laboratory (MSL)

landing system in 2012 illustrating the difficulty in high

mass Martian landings.

To achieve humanity’s goals for Mars exploration,

significant technology development is required. Mass

reducing technologies are particularly critical in this ef-

fort. Not only does a larger mass require more fuel to

launch, but it also carries significantly more kinetic en-

ergy that must be reduced to near zero if the vehicle is

to land safely. For all previous Mars missions, this ki-

netic energy has been primarily dissipated through aer-

odynamic drag during the hypersonic phase, which is

limited by the size of the entry vehicle. In addition, Pre-

vious Mars missions have shown that the majority of the

vehicle’s kinetic energy is dissipated during the hyper-

sonic entry phase, about 92% in the case of Mars Path-

finder [2].

During this hypersonic entry phase, there exists a

highly heated, ionized flow around the entry vehicle.

This entry plasma is inherently conductive, and may

thus be influenced by electromagnetic fields, facilitating

magnetohydrodynamic flow interaction. This flow in-

teraction can be used to provide a Lorentz force in addi-

tion to aerodynamic drag, thereby enhancing decelera-

tion capability without necessarily increasing the size

and mass of the entry vehicle forebody.

Magnetohydrodyanmic flow interaction for high

speed aerospace applications has been studied since the

dawn of the space race, with early theoretical studies da-

ting back to the late fifties and early sixties [3][4]. At

that time, practical implementation was limited due to

the difficulties in generating high strength magnetic

fields and electrical energy storage. Since that time,

however, dramatic advances in superconductivity and

electrical energy storage have been made, warranting

additional study.

Modern numerical techniques have been applied to

solve for the full-field flow around an entry vehicle sub-

ject to an applied magnetic field by previous research-

ers, however these studies usually focused on single ve-

hicle configurations and flow conditions, or Earth entry

[5]. A proposal is made to extend this analysis for mul-

tiple Mars entry vehicles along an entire simulated entry

trajectory to assess the system impact of magnetohydro-

dynamically enhanced drag for planetary entry vehicles.

Methodology: Non-lifting entry trajectories are

simulated for each vehicle, alkali metal seeding level,

and applied magnetic field strength using a one-dimen-

sional equilibrium chemistry model, and an assessment

is made of the impact of the applied magnetic field and

augmented drag on the trajectory.

Results: Results include simulated trajectories with

and without magnetohydrodynamic interaction across

various parameters. Following the presentation and dis-

cussion of the results, the observations are summarized

and conclusions drawn. Finally, recommendations for

future work on the subject are made.

Acknowledgments: This work is funded through a

NASA Space Technology Research Fellowship, grant

number NNX13AL82H. In addition, the authors would

like to thank Dr. Robert Moses of NASA Langley Re-

search center for his advice towards the completion of

this work.

References:

[1] Braun, R.D.; Manning, R.M., "Mars exploration

entry, descent and landing challenges," Aerospace Con-

ference, 2006 IEEE , 4-11 March 2006.

[2]Spencer, D., Blanchard, R., Braun, R., Pathfinder

Entry, Descent, and Landing Reconstruction,” Journal

of Rockets, Vol. 36, No. 3, May-June 1999.

[3] Bush, W. B., “A Note on

Magnetohydrodynamic-Hypersonic Flow Past a Blunt

Body,” Journal of the Aerospace Sciences, vol. 26,

1959, pp. 536–537.

[4] Jarvinen, P. O., On The Use of

Magnetohydrodynamics During High Speed Re-Entry,

Everett, Massachusetts: 1964.

[5] Fujino, T., Yoshino, T., and Ishikawa, M.,

“Numerical Analysis of Reentry Trajectory Coupled

with Magnetohydrodynamics Flow Control,” Journal of

Spacecraft and Rockets, vol. 45, 2008, pp. 911–920.



A DECELERATOR FOR PLUTO ENTRY, DESCENT, AND LANDING. B.D. Goldman1, K.T. Nock1, A.D.

McRonald1, J. Lin2, and S.B. Porter3, 1Global Aerospace Corporation, 12981 Ramona Blvd. Ste. E, Irwindale, CA,

91706 ([email protected]), 2ILC Dover LP, 1 Moonwalker Road Frederica, DE, 19946, 3Southwest Research Institute, 1050 Walnut St, Suite 300, Boulder, CO, 80302.

Introduction: The mission is to deliver a small

lander to the surface of Pluto, a planet with an atmos-

pheric pressure nearly 12 millionths of that on Earth

[1]. To add to the challenge, consider the scenario of

direct entry from a hyperbolic trajectory without orbit

capture or any form of propulsion. This spacecraft

could be traveling in excess of 14 km/s prior to en-

countering Pluto’s atmosphere. Imagine being able to

decelerate to a relatively more gentle velocity of 40 m/s

at the surface, using aerodynamic drag alone. We have

developed a planetary decelerator concept that would

enable propulsion-less entry, descent, and landing even

on planets with very thin atmospheres like Pluto. With

our unique entry technology, a lander could be deliv-

ered to the surface without retrorockets or parachutes.

While this particular mission is focused on Pluto, simi-

lar or spinoff concepts could be applicable on Mars

and Earth as well as natural satellites like Titan, Triton,

and possibly Io.

This mission follows in the footsteps of NASA’s

New Horizons flyby mission by conducting surface

exploration of Pluto. There is significant interest from

some members in the scientific community, including

those associated with New Horizons, in developing a

lander for Pluto [2][3]. Kuiper belt objects like Pluto

can provide a great deal of insight into the evolution of

our solar system, especially if there is the ability to

examine the surface directly. A lander could validate

measurements made by New Horizons in addition to

collecting new data from the surface.

References: [1] Sicardy, B., et al. (2016). Pluto's

atmosphere from the 29 June 2015 ground-based stellar

occultation at the time of the New Horizons flyby.

arXiv preprint arXiv:1601.05672. [2] Deutsch, L. J.,

Salvo, C., & Woerner, D. (2000). NASA's X2000 pro-

gram—An institutional approach to enabling smaller

spacecraft. Acta Astronautica, 46(2-6), 229-232. [3]

Osborne, H. (2015). New Horizons: Pluto lander would

be 'spectacular' next mission. International Business

Times Retrieved November 16, 2015, from

http://www.ibtimes.co.uk/new-horizons-pluto-lander-

would-be-spectacular-next-mission-1495456.



Combining Camera-based Hazard Detection and Terrain Relative Navigation in a SLAM-like approach. S. Woicke1 and E. Mooij2, 1Phd student, Faculty of Aerospace Engineering, Delft University of Technology, Klyverweg 1, 2629HS Delft, The Netherlands, [email protected], 2Assistant Professor, Faculty of Aerospace En-gineering, Delft University of Technology.

For future landing missions, touching down in

unknown or known, but unsafe environments might be necessary. An example for these kinds of scenari-os is landing on Europa: a body that is currently not mapped accurately enough to detect inherently safe landing regions. Landing in such an environment requires two capabilities: 1) Precision landing with terrain relative navigation (TRN) and a suitable guid-ance law, and 2) hazard detection and avoidance (HDA).

HDA creates a map of the body’s surface to select a safe landing site, at the same time TRN tries to lo-calize the vehicle based on images or a map. In robot-ics it is already common practice to combine locali-zation and enviroment mapping in one single algo-rithm – so-called simultaneous localization and map-ping (SLAM) algorithms.

An algorithm has been designed and developed

that combines TRN and HDA capabilities. Opposed to traditional TRN algorithms, this methods will not only enable more accurate state prediction, but will also make an attempt at predicting and propagating hazard maps and thereby trying to reduce the inaccu-racies in these maps. Figure 1 shows the simplified structure of the algorithm. It is obviously not possible to include the entire map in the filter state; therefore, multiple approaches to combining the map and the state are investigated.

Figure 1: Simplified Algorithm Structure

The hazard maps used for this algorithm will be

obtained using stereo vision [1, 2]. Thus, a camera- based, i.e., passive hazard detection algorithm is used. Since stereo vision measures the distance from the cameras to the imaged objects, these measure-ments can directly be used for TRN. Figure 2 shows a map as created by the stereo-vision algorithm.

Figure 2: Digital elevation model as obtained from

stereo vision algorithm

This paper discusses the first phase of this study, mainly focusing on using the HDA output as TRN measurements. Improving the hazard maps them-selves using prediction methods will be investigated at a later stage of the project. The scope of this paper is to present first preliminary results of this part of the algorithm and to discuss its applicability for fu-ture missions. Furthermore, an outlook to future work is given.

References: [1] Woicke S. and Mooij E. (2014) A Stereo-Vision Based Hazard-Detection Algorithm for Future Planetary Landers, LPI-001795 Ab-tract#8038. [2] Woicke S. and Mooij E. (2016) A stereo-vision hazard-detection algorithm to increase planetary lander autonomy, Acta Astronautica, Vol-ume 122, pp. 42-62, ISSN 0094-5765.



Combining Camera-based Hazard Detection and Terrain Relative Navigation in a SLAM-like approach. S. Woicke1 and E. Mooij2, 1Phd student, Faculty of Aerospace Engineering, Delft University of Technology, Klyverweg 1, 2629HS Delft, The Netherlands, [email protected], 2Assistant Professor, Faculty of Aerospace En-gineering, Delft University of Technology.

For future landing missions, touching down in

unknown or known, but unsafe environments might be necessary. An example for these kinds of scenari-os is landing on Europa: a body that is currently not mapped accurately enough to detect inherently safe landing regions. Landing in such an environment requires two capabilities: 1) Precision landing with terrain relative navigation (TRN) and a suitable guid-ance law, and 2) hazard detection and avoidance (HDA).

HDA creates a map of the body’s surface to select a safe landing site, at the same time TRN tries to lo-calize the vehicle based on images or a map. In robot-ics it is already common practice to combine locali-zation and enviroment mapping in one single algo-rithm – so-called simultaneous localization and map-ping (SLAM) algorithms.

An algorithm has been designed and developed

that combines TRN and HDA capabilities. Opposed to traditional TRN algorithms, this methods will not only enable more accurate state prediction, but will also make an attempt at predicting and propagating hazard maps and thereby trying to reduce the inaccu-racies in these maps. Figure 1 shows the simplified structure of the algorithm. It is obviously not possible to include the entire map in the filter state; therefore, multiple approaches to combining the map and the state are investigated.

Figure 1: Simplified Algorithm Structure

The hazard maps used for this algorithm will be

obtained using stereo vision [1, 2]. Thus, a camera- based, i.e., passive hazard detection algorithm is used. Since stereo vision measures the distance from the cameras to the imaged objects, these measure-ments can directly be used for TRN. Figure 2 shows a map as created by the stereo-vision algorithm.

Figure 2: Digital elevation model as obtained from

stereo vision algorithm

This paper discusses the first phase of this study, mainly focusing on using the HDA output as TRN measurements. Improving the hazard maps them-selves using prediction methods will be investigated at a later stage of the project. The scope of this paper is to present first preliminary results of this part of the algorithm and to discuss its applicability for fu-ture missions. Furthermore, an outlook to future work is given.

References: [1] Woicke S. and Mooij E. (2014) A Stereo-Vision Based Hazard-Detection Algorithm for Future Planetary Landers, LPI-001795 Ab-tract#8038. [2] Woicke S. and Mooij E. (2016) A stereo-vision hazard-detection algorithm to increase planetary lander autonomy, Acta Astronautica, Vol-ume 122, pp. 42-62, ISSN 0094-5765.



INFLATABLE DECELERATORS FOR HUMAN-SCALE MISSIONS TO MARS. R. A. Dillman1, S. J. Hughes1, J. M. DiNonno1, F. M. Cheatwood1, and R. W. Powell2; 1NASA Langley Research Center, 2Analytical Mechanics Associates Inc.

Abstract: Hypersonic Inflatable Aerodynamic De-

celerator (HIAD) technology is a leading candidate for delivery of large payloads to the surface of Mars in support of human-scale missions. A portion of NASA’s Evolvable Mars Campaign effort is directed to evaluation of the Entry, Descent, and Landing vehi-cles that will protect payloads during aerocapture into Mars orbit, and from orbit to the planet’s surface. These vehicles must protect the payloads from intense thermal environments while limiting deceleration loads and providing the necessary flight path control to stay on the desired trajectory despite variations in atmos-pheric conditions. HIAD technology allows a large-diameter entry vehicle to stow in a compact volume inside the launch vehicle shroud at Earth and to deploy to full size shortly before being needed at Mars. This paper will describe the baseline reference mission and payloads, the associated HIAD configurations, and the HIAD systems and design features that will allow the vehicles to decelerate from aerocapture and entry speeds, shield the payloads from the flight environ-ment, maneuver to reach the desired landing sites, and retract around the base of the lander after touchdown for ease of mission operations.



Lifting Entry Validation Experiment (LEVE) for Lifting Entry Atmospheric Flight (LEAF) Systems. G. Lee1, D. Sokol1 and F. Ross3, 1,2,3Northrop Grumman Aerospace Systems 1 Space Park, Redondo Beach, CA 90278 [email protected], [email protected], [email protected]

Introduction: Northrop Grumman, in collabora-

tion with L’Garde, has been developing the Lifting Entry Atmsopheric Flight (LEAF) system. The LEAF system is a lifting, hypersonic entry, maneuverable platform capable of performing long-term (months to a year) in situ and remote measurements at any solar system body that possesses an atmosphere. The Venus instantiation is called Venus Atmospheric Maneuvera-ble Platform (VAMP)[1] and the Titan version, T-LEAF[2].

LEAF’s ultra low ballistic coefficient combined with a lifting entry presents an innovative way to enter the atmosphere from space. Benefits of LEAF’s entry are (1) large payload capacity achieved by re-allocating the mass made available by the elimination of a heavy aeroshell, (2) reduced mission risk and crit-ical events by fully deploying in space and gently en-tering the atmosphere, and (3) benign entry loads (both thermal and structural) that allow use of a lightweight, flexible thermal protection system and readily availa-ble avionics and instruments.

Earth atmosphere presents an ideal environment in which to validate the lifting entry concept. The Lifting Entry Validation Experiment (LEVE) will be per-formed with an 8.5 meter wingspan vehicle. Several options are being considered to place the stowed LEVE vehicle in space, including sounding rockets, secondary payload launch manifests and SpaceX’s Dragon trunk. Additionally, the International Space Station may also be a suitable platform from which the LEVE vehicle could be released. Specific launch op-portunities will be studied and finalized as the LEVE vehicle is further matured. Similar to the LEAF system concept, the vehicle will be deployed exo-

atmospherically, prior to re-entry. To assure success, a series of Ground Deployment Demonstrations (GDDs) are planned prior to the LEVE earth entry flight demonstration.

Table 1 contains initial entry conditions for the LEVE vehicle’s re-entry to Earth atmosphere. Table 1 also contains the estimated entry loads as well as a comparison to Venus entry loads. As shown, it is ex-pected that the LEVE entry loads will exceed those at Venus (with VAMP), thereby providing a representa-tive demonstration of lifting entry.

Notional LEVE Re-entry condi-tions

LEVE Re-entry Loads

VAMP Entry Loads

v = 7.33 km/s Qdot = 30 W/cm2 Qdot = 15 W/cm2

h = 152 km Tmax stag = 1,796 K Tmax stag = 1,319K

γ = 1.4o gmax = 2.9 g gmax = 2.5 g

Table 1 LEVE Entry Conditions and Loads

In this presentation, we discuss the plans for LEVE, including the Ground Deployment Demonstra-tion of the 8.5 m wingspan vehicle. Additionally, we discuss the overall LEVE concept of operations and the aerothermal design.

References: [1] G. Lee, F. Ross, et al, “Venus Atmospheric

Maneuverable Platform (VAMP) Low Cost Pathfinder Mission Concept,” 11th Low Cost Planetary Missions Conference, Berlin, Germany, 9-11 June 2015; “Venus Atmospheric Maneuverable Platform (VAMP) Science Vehicle Concept,” Venus Science Priorities and In-strument Definition Workshop, Hampton, VA, 7-8 April 2015; “Venus Atmospheric Maneuverable Plat-form (VAMP) Air Vehicle Concept and Entry Concept of Operations,” 46th Lunar and Planetary Science Conference, Woodlands, TX, 16-20 March 2015

[2] F. Ross, G. Lee, et al, “Lifting Entry and At-mospheric Flight (LEAF) Applications at Various So-lar System Bodies,” 12th International Planetary Probe Workshop, Koln, Germany, 15-19 June 2015






Magnetohydrodynamic Energy Generation and Atmospheric Breathing Supersonic Retropropulsion for Mars Descent - POSTER. K. C. Gonyea1, H. K. Ali1, and R. D. Braun1. 1Georgia Institute of Technology, School of Aerospace Engineering. Northe Ave, Atlanta, GA, 30332.

Introduction: Design and analysis is performed on

the combination of an atmospheric breathing propul-sion system and magnetohydrodynamic energy genera-tion to land large-scale spacecraft on Mars. During the hypersonic phase of entry, kinetic energy from the space craft is transferred to the atmosphere through deceleration, creating a heated ionized, gas in the shock layer. Electrical energy is extracted and stored from this gas through magnetohydrodynamic energy generation. During descent, atmospheric carbon diox-ide is ingested via inlets on the entry vehicle. The in-gested carbon dioxide is used as the oxidizer for the retropropulsion system, combusting with onboard magnesium. The energy generated from the magneto-hydrodynamic energy generation is used to power a compressor, responsible for pressurizing the atmos-pheric oxidizer prior to combustion. This pressuriza-tion allows the propulsion system to operate at a higher ISP thus improving the engine performance. The sys-tem-wide effects of the magnetohydrodynamic energy generation system are analyzed including the available power production throughout the full trajectory as well as the implications of increased mass and complexity. These performance characteristics are compared to non-in-situ power generation methods. Two configura-tions were considered; the first, using the generated electrical energy to compress the oxidizer for use dur-ing steady-state thrusting and the second, using gener-ated electrical energy to compress stored atmospheric oxidizer for use during a terminal burn. For both con-figurations, flowpath design is discussed to allow for the operation of each system. In addition, compressor design is analyzed to ensure that sufficient pressure ratios are attained.



Mars 2020 Terrain Relative Navigation Accommodation Aaron Stehura1, Paul Brugarolas1, Allen Chen1, Andrew Johnson1, Swati Mohan1, Jim Montgomery1 1Jet Propulsion Laboratory, California Institute of Technology, 4800 Oak Grove Drive, Pasadena, CA 91109

Abstract: The Terrain Relative Navigation (TRN) system is

an enabling Entry, Descent, and Landing (EDL) technology slated for inclusion in the Mars 2020 mission [1]. TRN provides real-time, autono-mous, terrain-relative position determination and gen-erates a landing target based on a priori knowledge of hazards. TRN is composed of the Lander Vision Sys-tem (LVS) and the Safe Target Selection (STS) algo-rithm. The LVS generates a map-relative localization solution by fusing measurements from a visible-wavelength camera and an inertial measurement unit using the Map Relative Localization algo-rithm operating on a high-performance compute ele-ment. Updated state knowledge is provided to the spacecraft navigation filter, which uses the STS algorithm to direct a divert maneuver away from known hazards within an onboard map.

Previous papers have discussed the prototype LVS that was developed and demonstrated over relevant terrain [2], the architecture and capability of the STS algorithm [3], and the analyses that show how TRN enables access to new, scientifically compelling land-ing sites [2]. Now that TRN is officially part of Mars 2020's baseline configuration, this paper shows how the system is integrated into the existing EDL architec-ture. The constrained set of deltas to the EDL system and the overall Mars 2020 accommodation story are presented. This example leads to a method in which the LVS can be employed as a bolt-on sensor when seeking to avoid hazards while landing on other bod-ies. References:

[1] Allen Chen et al. (2015) 2015 Update: Mars

2020 Entry, Descent, and Landing System Overview, IPPW12 Presentation #2104.

[2] Aaron Stehura et al. (2015) The Future of Land-ing: Terrain Relative Navigation From Prototype to Mars 2020, IPPW12 Presentation #3104.

[3] Paul Brugarolas et al. (2015) On-Board Ter-rain Relative Guidance-Target Selection for the Mars 2020 Mission, IPPW12 Presentation #3105.



SENSOR DATA FUSION FOR HAZARD MAPPING AND PILOTING. K. Kanani1, R. Brochard1, F. Hen-nart1, A. Pollini2, P. Sturm3, O. Dubois-Matra4, S. Vijendran4, 1Airbus Defence & Space, 31 rue des cosmonautes, 31402 Toulouse Cedex, France ( [email protected], [email protected], [email protected]), 2CSEM, Rue Jaquet-Droz 1, CH-2002 Neuchâtel ([email protected]), 3INRIA Grenoble Rhône-Alpes, Inovallée, 655 Avenue de l’Europe, Montbonnot, 38334 Saint Ismier Cedex, France ([email protected]), 4ESA-ESTEC, Keplerlaan 1, 2201 AZ Noordwijk, The Netherlands ([email protected], [email protected]).

Introduction: Autonomous landing on Mars,

Moon or asteroids may require an autonomous Hazard Detection and Avoidance (HDA) system. Past studies on HDA or actual missions (such as Chang’E 3) dealt with the use of camera or LiDARs separately to detect dangerous slopes, boulders and shadowed areas. The present work, performed in the frame of an ESA Tech-nology Research Program, proposes to use jointly a camera and a LiDAR to take advantage of each while mitigating their drawbacks, consequently improving the HDA performances. Various algorithmic solutions and sensor configurations are proposed and tested in the Mars and asteroid landing cases.

Challenge of slope mapping: In the present work, we focus on the computation of the slope map, which is deemed the most challenging part of HDA. The oth-er components of the GNC and HDA system (rough-ness and shadow, decision making...) are mainly re-used from previous studies, even if the relative naviga-tion has been improved with altitude measurements from the LiDAR to meet the requirements of slope mapping.

The joint use of LiDAR and camera data enables to release the requirements on the LiDAR, especially for Mars landing, in terms of mass, power, resolution and frame rate, in comparison with the use of a LiDAR alone. The selected LiDAR is a last-generation flash LiDAR built by CSEM [2], whereas the chosen camera is based on heritage from ESA and NASA studies. The retained algorithmic solution for slope mapping is composed of (1) Shape from motion (3D reconstruc-tion) [2] and (2) Machine learning.

HDA framework: Given the limited optical power of the LiDAR light source and the high altitude at which hazard mapping has to be performed, LiDAR data is very noisy and barely usable as it. Our HDA solution starts with a temporal and spatial denoising of the LiDAR signal. The 3D reconstruction is then per-formed using LiDAR and camera measurements. A classification step, based on machine learning, is then applied, using the 3D reconstruction output and the camera images, to provide a improved slope map. The slope map is finally combined with shadow and roughness maps, using a fuzzy logic technique to se-

lect a safe landing site, with an observed false positive rate less than 1%.

Evaluation with a high-fidelity GNC simulator: The HDA system is assessed within a complete GNC simulator, with image simulation and image processing in the loop. It includes realistic sensor data simulation, taking into account a model of the flash LiDAR devel-oped by CSEM, and considering real Mars and aster-oid terrains, extracted from NASA’s PDS database. This GNC framework enabled Monte Carlo analyses, to assess the end-to-end performances of the solution. The performances of our HDA system, as well as an analysis on on-board implementation, will be de-scribed in this presentation.

References:

[1] “Multi-view 3d reconstruction of asteroids”. A. Delaunoy, K. Kanani, P. Sturm, O. Dubois-Matra. 5th International Conference on Astrodynamics Tools and Techniques. May 2012. [2] “Flash optical sensors for guidance, navigation and control systems”. Alexandre Pollini, 35th Annual AAS guidance and control conference, Feb. 2012.

Example of considered sensor configuration: the Li-

DAR and the camera have the same resolution but the LiDAR has a narrow FoV.

Example of 3D reconstruction: (left) with realistic nav-igation errors, (right) with perfect navigation, (middle)

ground truth.












TECHNOLOGY OVERVIEW AND ASSESSMENT FOR SMALL-SCALE EDL SYSTEMS. C. R. Heidrich1 and R. D. Braun2, 1School of Aerospace Engineering, Georgia Institute of Technology, Atlanta, GA 30332, 2School of Aerospace Engineering, Georgia Institute of Technology, Atlanta, GA 30332.

Introduction: Technologies to enable high-mas

atmospheric entries are a current trend in the entry, descent, and landing (EDL) research community. This is motivated by missions to land large rovers or hu-mans at Mars and other bodies. EDL systems for low-mass payloads, in contrast, have attracted less attention in current research. Significant potential exists in ap-plications of small-scale EDL systems. These architec-tures can provide a simple and efficient means of achieving scientific and mission support objectives.

The movement toward flying smaller spacecraft emerged out of academia through strict enforcement of mass and volume limitations and by defining standard mechanical and electrical interfaces. In recent years, NASA has begun to actively integrate small spacecraft into its mission portfolio and is reaping the rewards. The value proposition posed by small spacecraft is high due to their dramatically lower cost and more forgiving risk posture compared with large, “traditional” space-craft with payload-driven requirements. The vision of the current research is to apply the tenets of the small spacecraft movement toward an EDL system for sec-ondary payload missions. An EDL capability for small spacecraft would enable scientists to recover payloads from Earth orbit as well as broaden the reach of small spacecraft to landing on other celestial bodies with substantial atmospheres (e.g. Mars, Venus, and Titan).

Several challenges exist with future development of small-scale EDL systems. Multiple modes of failure are inherent in atmospheric entry missions due to large performance uncertainties. Volume and space limita-tions may dictate redesign, cross-utilization, or com-plete removal of systems to achieve feasibility. Recon-ceptualization of heritage EDL architectures creates a higher susceptibility to failure. Removal of redundant systems decreases robustness of the system to unex-pected variables in flight. For example, active control and stabilization during reentry may be impossible for small-scale vehicles; a passive or semi-passive EDL timeline may be the only option. Ballistic spin-up prior to entry was flown on Mars Pathfinder (MPF) and Mars Exploration Rover (MER) to provide entry atti-tude control. [1] Use of ballistic spin-up with small EDL systems may impose unacceptable complexity on the parent mission. An important drive for this research is to identify the trade between system design and vehi-cle flight performance. Each system is subject to con-straints on EDL timeline, landing g’s, and packaging volume. It is important to quantify how the scaling of

each system will affect trajectory dispersions and risk expectancies.

This research will explore the design space for small-scale EDL systems. This design space is unique as it includes technologies often considered infeasible on large scales. The volumetric efficiency of a system (performance per unit volume) will heavily influence its applicability on small scales. Performance of entry systems with atypical placement and orientation in the vehicle will also be considered.

Passive design architectures, particularly for de-scent and landing systems, will provide the most diffi-cult scaling challenges. This is due to the fact that many technology options require static hardware that scales inefficiently. For example, parachutes requiring drogue or mortar deployment will package less effi-ciently. Crushable impact systems require large longi-tudinal space in the aeroshell, and only a portion of the crushable mass effectively absorbs energy. It is im-portant to study systems that package well and provide the greatest performance per volume on a small scale.

Technology overview. Several options have been studied for high-mass EDL concepts. [2] The goal of this research is to study applications of these concepts on small scales. Based on the history of small entries at Earth and Mars, typical outer mold lines (OMLs) are rigid blunt aeroshells. This is typically due to the herit-age data and flights of sphere-cone geometries. How-ever, novel concepts such as deployable or inflatable hypersonic decelerators may prove viable when subject to volume constraints. [3]

Descent systems provide previously under-utilized technology concepts. Past small architectures have included a disk gap band (DGB) or similar parachute. These systems typically have a narrow flight envelope for deployment. Parachutes also impose a moderate dry mass to stow and deploy the canopy. Novel technology, including the auto-rotor, propulsive, or drag ribbon concepts, may present options for increasing volume efficiency of these systems.

Landing systems present a significant obstacle for volume reduction in small EDL architectures. Of par-ticular interest are largely passive landing systems, as they provide reliability without requiring extraneous sensor feedback near the surface. These include legs, crushable structures, shell landers, and penetrators. A single element unique to each approach is the use of a deformable impact structure. In each concept, the ki-netic energy of the vehicle is absorbed by the vehicle



structure. A majority of the impact stroke is limited by the OML forebody geometry. Localized stresses in the material could potentially lead to material failure dur-ing impact; the residual energy would then be transmit-ted directly to the payload. [4] Due to the uncertain nature of impact attenuation on small scales, it may become necessary in this research to distinguish soft landers and penetrating probes.

In order to overcome challenges associated with downscaled EDL missions, it is useful to formulate atypical architectures. For the descent stage, this in-cludes developing compact parachute packaging tech-niques. A flattened or annular canister requires less longitudinal space in the vehicle. Largely untested con-cepts present viable options for efficient volume utili-zation of the descent stage. In some cases, it may be necessary to completely eliminate the descent stage, with the landing system absorbing the remaining kinet-ic energy on impact.

Potential cross-utilization of EDL systems also cre-ates opportunities for volume optimization in the vehi-cle. A forebody decelerator, such as HIAD or nano-ADEPT, may be designed to act as a secondary impact attenuator on landing. A parachute or drag ribbon may extricate the payload from the aeroshell prior to land-ing, significantly reducing the landing stress on the payload. These novel concepts, typically unacceptable for a high-mass architecture, present unique solutions to space limitations on small scales. The approach to take measured risk with small EDL systems makes ap-plication of low technology readiness level (TRL) sys-tems more viable.

Example mission design. It is useful to develop a technology baseline for small entry systems. An exam-ple mission design was performed to identify the mini-mum configuration to provide a rough landing (surviv-able for a sensitive payload) at Mars. The mission pa-rameters for a hypothetical small Mars lander, framed as a secondary payload, are given as: (1) The payload mass and volume are 3kg and 3U, respectively (2) Re-strictions on the primary vehicle limit the capsule height to 0.8 m (3) The payload shall experience less than 12 g's sustained (4) The bondline temperature shall not exceed 250 °C due to aerothermal heating (5) The payload shall experience less than a 50 g landing impulse. A potential packaging configuration for this mission is outlined in Figure 1. Figure 2 illustrates the subsystem sizing, where the total entry mass was 23 kg.

Figure 1. Packaging configuration for a small passive Mars lander.

Figure 2. Subsystem sizing for a small passive Mars lander.

Maturation of small EDL technologies greatly in-creases our aptitude to return from space. Non-conventional science and exploration missions are like-ly to achieve fruition on a small-scale, affordable plat-form. The potential to piggyback existing missions as a passive secondary will push these possibilities closer to realization. This research outlines the potential design space for small EDL systems. Scalability and volume packaging constraints present a challenge for feasibil-ity. Minimum mass configurations may be infeasible with current technologies due to packaging constraints. Future development of this subject will require theoret-ical and experimental testing of small-scale novel tech-nologies. References: [1] Braun, R. D. and Manning, R. M. (2006) IEEE, 18–pp. [2] Peacocke, Lisa, et al (2011) IPPW. [3] Smith, Brandon, et al. (2014), IPPW, Vol. 11. [4] Jones, R. H. (1970), NASA-SP-8046.



Development of Atmospheric Entry Trajectory and Aerothermodynamics Code & Design of Deployable Aer-

odynamic Decelerator for Sample Return Mission A. Jasjukevics1 and G. Markelov

2,

1ATG Europe\University of

Glasgow (Huygensstraat 34, 2201 DK Noordwijk, Netherlands, [email protected]), [email protected]

Abstract

Deployable and Inflatable Aerodynamic Decelera-

tors for atmospheric entry vehicles are a promising

concept that has been discussed since the dawn of

space age, which has lately become feasible due to

introduction of new structural and thermal protection

materials. It has been decided to study the concept, find

a possible application for it and develop a preliminary

design of a Deployable Aerodynamic Decelerator

along with programming a design tool.

Engineering high speed aerodynamics model for

blunt bodies has been programmed and validated. It

allows inputting a capsule geometry directly to the

code without having to generate aerodynamic database

with complex and computationally expensive methods.

Based on it, a 3 Degrees of Freedom trajectory code

has been developed and validated. Correlation based

heating models have also been implemented. The result

is a tool that was used in the next part of the project to

quickly determine aerothermodynamics of design itera-

tions.

A tension cone structure, protected by a flexible ab-

lator and deployable using Shape Memory Alloy, has

been designed as an alternative for Earth Re-entry Cap-

sule for European Space Agency's Phobos Sample Re-

turn mission. Aerothermodynamic, stability, structural,

trajectory and thermal analyses have been performed as

a part of preliminary design.

Future work has been outlined for the further de-

sign stages. Alternative applications to proposed con-

cept have been suggested.



Gas Generators and their potential to support Human-Scale HIADs (Hypersonic Inflatable Aerodynamic De-

celerators) Richard J. Bodkin1, F.M. Cheatwoodr2, Robert A. Dillmanr3 ,Johm M. DiNonno4 Stephen J. Hughes5

Melvin H. Lucy6 1NASA Langley Research Center, MS 432 Hampton, VA 23681, USA; Email: [email protected] 2NASA Langley Research Center, MS 489 Hampton, VA 23681, USA; Email: [email protected] 3 NASA Langley Research Center, MS 489 Hampton, VA 23681, USA; Email: [email protected] 4NASA Langley Research Center, MS 432 Hampton, VA 23681, USA; Email: [email protected]

5NASA Langley Research Center, MS 432 Hampton, VA 23681, USA; Email: [email protected]

6NASA Langley Research Center, MS 488 Hampton, VA 23681, USA; Email: [email protected]

.

.

Abstract: Hypersonic Inflatable Aerodynamic Decel-

erators (HIADs) are a technology allowing deployable

protective aeroshell larger than will fit inside a launch

vehicle shroud to be used for future large or heavy

payloads. The technology is progressing from the

research/experimental stage towards flight applications.

The previous experimental systems have utilized blow-

down inflation systems (Figure 1 and 2). As the tech-

nology progresses from the 3-m diameter experimental

scale to the 20-m operational scale the mass penalties

of carrying compressed gas has led the HIAD-2 ground

team to start a research effort into gas generator

technologies. This paper will present the results from a

survey on the current state of gas generators, and

discuss the pros and cons of the various technologies

available to support large scale HIADs.

Figure 1: IRVE-II Inflation System

Figure 2: IRVE-3 Inflation System



The Horizon 2020 project IRENA (International Re-Entry demoNstrator Action) Jean-Marc Bouilly 1, Aurélien Pisseloup 1, Pierre W. Bousquet 2, Hendrik Weihs 3,Cosimo Chiarelli4, Martins Sudars4, Georgios Vekinis 5, Tetsuya Yamada 6, Ethiraj Venkatapathy 7,and the IRENA Team

1 Airbus Defence and Space - Bordeaux, France 2 CNES – Toulouse, France, 3 DLR - Stuttgart Germany, 4 Thales Alenia Space – Torino, Italy, 5 NCSR “Demokritos” – Athens, Greece,

6 JAXA - Sagamihara, Japan, 7 NASA Ames Research Center– Mountain View, CA,USA, Main author contact : Airbus Defence and Space – Space Systems BP 20011, av. du general Niox, 33 165 Saint-Médard-en-Jalles Cedex, France email : [email protected]

Abstract: In January 2014, the ISEF participants insisted on the importance of “fostering international cooperation for additional space exploration projects”. The need for demonstrators in atmosphere entry/re-entry and the potential for international cooperation in this area were recognized by the IRENA partners. The project was selected out of the 2014 Space Call COMPET-09-2014: “Technology “demonstrator pro-jects for exploration” from the European Union’s Horizon 2020 research and innovation programme.

This action received funding under grant agree-ment No 640277. It started in January 2015, and was completed in April 2016, after addressing two main objectives :

- Define two technology demonstrator projects to validate advanced entry/re-entry technolo-gies for further space exploration missions

- Create the ground for their actual implementa-tion (funding, governance, partnership) in an international cooperation framework.

To achieve these objectives, IRENA relied on an international and complementary team: four major European and international space agencies involved in ISEF and ISECG (CNES, DLR, NASA, JAXA), the two European industry leaders in entry/re-entry and space exploration (Airbus Defence and Space, Thales Alenia Space) and a research institute also in charge of dissemination and exploitation (Demokritos).To max-imise the chances of the projected demonstrators mate-rialising in the future, the international dimension is essential, which explains why two key non-European actors (JAXA and NASA) have been invited to join a team based otherwise on the most prominent European actors in the field.

In addition, dissemination / communication work-shops were organized throughout the whole project duration to enlarge beyond the project team with stakeholders not directly involved in the consortium, such as ESA.

After a quick reminder of IRENA’s organisation and workflow, the presentation/paper will provide an overview of the main outcomes of the 3 steps of the project:

- Review of demonstration needs for entry/re-entry between Europe, the USA and Japan: This resulted in identifying the EDL technol-ogies requiring increase of the TRL level in near future, and outlining possible flight and ground demonstration missions to cover the technological gap.

- Selection of demonstrator concepts, which led to focus on deployable decelerators, aerocap-ture and TPS ground validation aspects: thanks to a series of dedicated workshops, a phase 0 development stage was achieved, also allowing to estimate a preliminary cost.

- Selection the 2 most promising demonstrator projects (Deployable Aerodynamic Decelera-tor and Earth Aerocapture Demonstrator), and outline of their possible implementation plan in terms of governance, funding and interna-tional cooperation.

The IPPW13 workshop represents a great oppor-

tunity for the IRENA team to promote the outcomes of the project and to receive valuable feedback from the community.

More online at www.irena-project.eu



AIR DUST REMOVAL TOOL (AirDRT): A NOVEL CONCEPT FOR CLEANING ROCKS AND

SPACECRAFT SURFACES OFF DUST ON MARS USING HIGH SPEED PROPELLER. K. Zacny1, S. In-

dyk1, Z. Fitzgerald

1, J. Spring

1, P. Chu

1, G. Paulsen

1, Marcus Traylor

1,

1Honeybee Robotics (398 W Washington

Blvd, Pasadena, CA 91103, [email protected].)

Introduction: Exploration of Mars requires access

to pristine rock surfaces. The two rovers currently op-

erating on Mars have two different approaches to per-

forming rock surface cleaning functions (Figure 1). The

first approach uses the Rock Abrasion Tool (RAT) on

the Mars Exploration Rovers Spirit and Opportunity

[1]. The RAT is a turret mounted tool capable of grind-

ing and/or brushing a rock surface at high speeds (up to

3000 rpm). During grinding, the grind bit removes the

rock surface while the brush simultaneously clears the

abraded cuttings. To brush without grinding, the grind

bit retreats from the rock surface allowing only the

brush to contact and clear the rock surface.

The second approach to rock cleaning is used on

Mars Curiosity Rover. Dust Removal Tool or DRT [2]

is also a turret mounted tool with a goal to brush away

surface dust using high speed (500 rpm) twin brushes.

DRT does not have rock grinding capability.

Figure 1. Rock surface cleaning tools with RAT on

MER (left) and DRT Curiosity (right).

There are other ways that rock cleaning function

can be achieved. The above mentioned methods use

turret mounted tools with high speed stainless steel

brushes (high speed is needed to impart sufficient cen-

trifugal force to dust particles to move them out of the

‘clean’ area). Another method to rock surface cleaning

is to use a dedicated brushing bit (with stainless steel

brushes) that could be deployed by and powered by a

drilling system [3]. A prototype of such a tool was de-

veloped and successfully tested. The Rock Abrasion

and Brushing Bit (RABBit) is essentially a planetary

gearbox that emulates operation of the RAT, but is

driven by a drill (Figure 2).

In addition, non-contact means of removing dust

include a stand-alone high pressure canister (air duster)

or propeller driven tool. Air duster is relatively simple

approach but utilizes a consumable – high pressure gas.

Once the gas has been depleted, the tool is no longer

functional. It is possible to replenish gas by compress-

ing Martian air, though this would require an actuator

and compressor,increasing the system size and mass.

A high speed propeller approach could solve the

‘consumable’ problem and at the same time preserve

the numerous benefits of compressed air approach.

Figure 2. Rock Abrasion and Brushing Bit.

Air Dust Removal Tool (AirDRT): The AirDRT

could be a stand-alone, turret mounted tool (just like

RAT or DRT) or a drill deployable tool (just like

RABBit). There are advantages and disadvantages to

both options. The drill deployable option requires the

drill to mate/demate with the AirDRT each time the

tool needs to be used. This style AirDRT also needs

significant gear reduction ratio to achieve >20,000 rpm

propeller speed. Since the tool would be driven by the

drill itself, this wears the drill, reducing its life.

A stand alone, arm deployable AirDRT is a simpler

option unless there is no space on the turret. A proto-

type AirDRT fits within a 100 mm diameter and 100

mm long cylinder (i.e. shorter but large diameter than a

coke can) and weighs ~250 grams (Figure 3).

Figure 3. AirDRT.

It uses brushed DC motor and in turn requires just 2

wires for On/Off operation. An additional 4 wires

would be needed for a heater and a temperature sensor

to ensure the motor is at the right temperature for its

operation. The tool requires a duration of less than 1



minute to clear a surface of dust. The longevity of the

tool is based on the measured life of flight grade

brushed actuators at 16M cycles, allowing approximat-

ly 1000 deployments over the life of the tool [4].

Testing of AirDRT: A conventional ducted fan

(with blades designed for operation at 1 atm) was ini-

tially used in a Mars chamber to prove the viability of

this approach (Figure 4). A number of variables were

investigated including a distance from the surface, rpm,

angle wrt surface, shape of the nozzle (i.e. angle of the

divergent nozzle), and moving the AirDRT above the

surface in a sweeping motion (similar to the operation

of DRT on Curiosity). It was found that the ducted fan

worked best at approximately 1 cm off the surface and

at speeds in excess of 20,000 rpm (the exact speed was

difficult to measure). Most nozzle configurations creat-

ed a ‘dead’ spot in the middle – probably because air

flow was absent in that region. The tool was most ef-

fective in removing dusts when it was moved in a

sweeping motion above the surface. This requirement

could be eliminated through a better design of the con-

verging nozzle. The nozzle, for example, could pro-

duce vortices that might be more effective in cleaning

larger surface areas.

In all cases, the ducted fan succesfully cleaned rock

surfaces off dust in 1 minute or less and as such

demonstrated viability of this approach for future Mars

surface missions.

Deployment at 45 deg Deployment at 90 deg with

a long converging nozzle

Before After

Figure 4. AirDRT tests in a Mars chamber to re-

move abraided rock surface dust.

Benefits of AirDRT: The AirDRT requires only 6

wires: 2: Brushed DC motor, 2: Heater, 2: Temp sensor

(not included shielding, redundant wires etc.). It has

simple On/Off operation (no speed control needed). It

can be either preloaded against the surface or could be

swept by the robotic arm (as demonstrated with Curios-

ity’s DRT). The AirDRT produces no contamination

(brushes could contaminate surface if they breakoff),

does not use a consumable material (e.g. gas), requires

less than 1 min of operation (brushing operations on

MER and MSL takes ~20 minutes). In addition,

brushed actuator can provide for over 1000 x one

minitue long deployments. The AirDRT is also very

effective since it clears cracked surfaces where brushes

can’t penetrate and also does not scratch soft surfaces,

the way brushes do.

Optimizing AirDRT: As mentioned earlier, initial

proof of concept tests used a ducted fan designed for

operation at 1 atmospheric pressure. Hence significant

improvement in performance is expected by designing

the system for operation at a pressure 100x lower (i.e.

Mars pressure of approximately 10 mbar). The areas of

improvement include propeller shape and pitch angle,

diameter and shape of the ducted fan housing and con-

verging nozzle (e.g. vortex inducing nozzle), shape and

size of dust deflector plate (our tests to date have

shown such a plate might not actually be needed), an-

gle of operation wrt surface, arm sweep motion (might

not be needed).

Other uses of AirDRT: Although the predominant

purpose of the AirDRT would be to clear rock surfaces

of dust or remove ground up rock after abrading the

rock surface, there are other potential uses of AirDRT

on Mars. The AirDRT could be used to remove dust

from solar panels (if present), Radioisotope Thermal

Generators (RTG) radiator fins, and camera optical

lenses, and in turn make these subsystems more effi-

cient. The AirDRT could be used as a trenching device

to create holes in loose sediments and in turn could

help during the search of the Mars2020 core tubes

(other approaches such as scoops could potentially

damage the tubes). The AirDRT operated in reverse

could be used as an affective sample acquisition system

(a true Martian vacuum cleaner).

References: [1] Gorevan et al., (2003), Rock Abra-

sion Tool: Mars Exploration Rover Mission. JGR,

108., [2] Davis et al., (2012), Mars Science Laborato-

ry’s Dust Removal Tool, AMS, [3] Zacny et al. (2012),

Mars Drill for the Mars Sample Return Mission with a

Brushing and Abrading Bit, Regolith and Powder Bit,

Core PreView Bit and a Coring Bit, IEEE Aeroe Conf.,

[4] Phillips et al., (2012), Development of Brushed and

Brushless DC Motors for use in the ExoMars Drilling

and Sampling Mechanism, AMS.



Astrobiology at CNES : research and development programmes paving the way for the future P. Chazalnoël1

1Centre National d’Etudes Spatiales, 18 av E. Belin, 31400 Toulouse, France, [email protected],

Introduction: Research for life on other system solar

bodies and the study of the origin of life on Earth are essen-

tial thematic for planetary exploration. New instruments or

more specific and sensitive instruments need to be de-

veloped for future mission. They need also to be minia-

turized in order to reduce mass, volume, power con-

sumption. This talk will present some studies on-going

in CNES for future instruments for planetary mission

and also a miniaturized sample-holder to allow direct

analyses of a set of extraterrestrial

Research and development programmes at

CNES: Since 2006, the CNES emits every year a spe-

cific announcement for astrobiology research studies.

Various studies have been performed and others are

paving the way for future instruments. Some of them

will be presented.

Sample holder A miniaturized sample-holder [1]

has been designed and built to allow direct analyses of

a set of extraterrestrial grains confined in a sealed con-

tainer implementing three layers of containment and

remotely positioned in front of the X-ray or laser

beams of the various setups. The grains are held in

several thin walls (10 µm) ultrapure silica capillaries

which are sufficiently resistant for manual/remote-

controlled micro-manipulation but semitransparent for

the characteristic X-rays, Raman and IR radiations.

Miniaturized pressure/temperature sensors located in

each container periodically monitor the integrity of the

ensemble, ensuring leak proof conditions.

Life marker chip Biochips are promising instru-

ments for the search for organic molecules in planetary

environments. We have developed a biochip based on

antibodies and aptamers in order to detect the presence

of life marker in in-situ samples [2]. We studied the

different elements of the biochip: the slide, the ligands

and the chemical bonds. Resistances to space condi-

tions of the biochip have been performed [3].

Liquid chromatography The gas chromatography is

a very powerful instrument for the study of in-situ

sample. This technic has flown on different space mis-

sion. We are studying an additional new one, based on

liquid chromatography. A breadboard is being devel-

oped.

References:

[1] A. Simionovici and CNES, (2010), INPI patent

pending. [2] A. Le Postollec et al. (2007) SF2A hal-

00255886. [3] M. Baqué et al. (20 11) Astrobiology

volume 11, number3




CHARACTERIZATION OF A METHOD FOR INVERSE HEAT TRANSFER. M. E. Pizzo1, K. S. Bey2, and

D. E. Glass3, 1 Old Dominion University, Norfolk, VA 23529, USA, [email protected], (609) 517-7660, 2NASA Lang-

ley Research Center, Hampton, MS 190, VA 23681, USA, [email protected], and 3NASA Langley Research Cen-

ter, Hampton, MS 190, VA 23681, USA, [email protected].

Introduction: It is often impractical to instrument

the external surface of high speed vehicles. Temperature

can instead be measured internally and direct and in-

verse methods can be used to estimate temperature and

heat flux on the external surface. Internal temperature

data are measured using thermocouples embedded

through-the-thickness of a high temperature material

and are, often time, noisy. Two embedded thermocou-

ples are required to solve direct and inverse problems,

and filtering schemes are used to reduce noise in the

measured data.

Accuracy in the estimated surface temperature and

heat flux is dependent on several factors. Factors in-

clude the effectiveness of data filtering, the instrumen-

tation locations through-the-thickness of a material, the

sensitivity to depth inaccuracy, and the sensitivity to

thermocouple inaccuracy. The effect of these factors on

solution accuracy is studied.

A numerical study is performed to determine if there

is an optimal depth to embed one thermocouple

through-the-thickness of a material assuming that a sec-

ond thermocouple is embedded on the back-face. Solu-

tion accuracy will be discussed for each selected instru-

mentation location. Moreover, the sensitivity to both

depth and thermocouple inaccuracy are quantified using

numerical simulation, and results discussed. Finally, the

importance of performing a filtering study prior to solv-

ing the direct and inverse methods will be discussed.

Presentation Format: A general overview and the

numerical results will be presented during an oral

presentation. Mathematical details will be presented

during a poster presentation. The oral and poster presen-

tations, although separate, will complement each other.

The title and abstract remain the same for both.

References: [1] Michelle E. Pizzo, David E. Glass,

and Kim Bey. “Analysis of Internal Thermocouple Data

in Carbon/Carbon Using Inverse Heat Conduction

Methods,” 54th AIAA Aerospace Sciences Meeting,

AIAA SciTech, (AIAA 2016-0508).






MARSDROP: GETTING MINIATURE INSTRUMENTS TO THE SURFACE OF MARS AS SECONDARY

PAYLOADS. R. L. Staehle1, M. A. Eby

2, J. J. Lang

2, J. A. Lang

2, R. M. E. Williams

3, J. S. Boland

1, R. Castano

1,

M. S. Lane1, C. A. Lindensmith

1, S. Spangelo

1.

1Jet Propulsion Laboratory, California Institute of Technology (4800

Oak Grove Drive, Pasadena, California 91109, [email protected]), 2The Aerospace Corporation (2310 El

Segundo Blvd., El Segundo, CA 90245, [email protected]), 3Planetary Science Institute (1700 East Fort Low-

ell, Suite 106, Tucson, AZ 85719, [email protected])

Small (~1 kg) instrument payloads could be carried

to Mars’ surface utilizing the MarsDrop delivery sys-

tem, based on an architecture that takes advantage of

the extra cruise stage mass capabilitiy available on

most Mars missions[1].

From canyons to glaciers, from geology to astrobi-

ology, the amount of exciting surface science awaiting

us at Mars greatly outstrips the available mission op-

portunities. Whether from the destination risks or just

from the significant expense of a traditional Mars

lander, the majority of proposed scientific surface mis-

sions are eliminated from consideration. By utilizing

The Aerospace Corporation’s Reentry Breakup Re-

corder (REBR) entry system already proven at Earth

with entry velocities greater than for Mars missions,

and adding a parawing for descent and landing that has

been tested above the Earth’s stratosphere at Mars dy-

namic pressure and density, a 3 kg entry vehicle can

deliver small instruments to targeted locations. Such a

vehicle could be accommodated on direct-entry Mars

missions for <10 kg total mass allocation per Mars-

Drop secondary lander, including the attach-

ment/jettison equipment on the primary cruise stage.

Depending on orbital parameters, slightly larger mass

allocations could accommodate such secondary landers

on missions where an orbiter is the primary spacecraft.

CubeSat and smallsat-class componentry, such as

that utilized for JPL’s Interplanetary NanoSpacecraft

Pathfinder In a Relevant Environment (INSPIRE)[2],

Mars CubeSat One (MarCO)[3], and other sources,

would provide the needed electrical power, computing,

and telecommunications resources to enable surface

operations for 90 sols, and potentially much longer.

MarsDrop’s small size could enable sterilization of

its components, sterile assembly, and encapsulation in

a sterile plastic shrink-wrap bag for ground handling.

This bio-barrier bag would later burn off during hyper-

sonic Mars entry. As a result, “special regions” on

Mars, where the presence of part-time liquid water is

possible, could be feasible targets within NASA plane-

tary protection guidelines. Furthermore, sampling ma-

terial that might have rolled down the side of a crater

exhibiting potentially-wet recurring slope lineae

(RSLs) could be feasible, along with other targeted

destinations using terrain relative navigation and steer-

ing MarsDrop’s parawing.

A variety of miniaturized instruments are being de-

veloped for different Solar System destinations. These

include cameras, spectrometers, seismometers, and

Mars weather instruments. Among other instruments

being miniaturized is a digital holographic microscope

[4]. With its fraction-of-a-millimeter depth of field,

such a microscope could survey a significant volume of

water-diluted sediment, taking inventory of solid parti-

cle shapes, sizes, and distribution, while monitoring for

unexpected features such as non-Brownian movement

contrary to flow. A terrestrial prototype of this instru-

ment has observed living organisms coming out of just-

melted Greenland glacier ice. Utilizing hardware and

software similar to JPL’s CubeSat Onboard Processor

Validation Experiment (COVE) [5] and Autonomous

Exploration for Gathering Increased Science (AEGIS)

[6], the most relevant information may be extracted for

downlink within a modest >0.5 MB/day budget via

existing Mars orbiting relay assets.

References: [1] Staehle R. L. at al. (2015) “Multi-

plying Mars Lander Opportunities with MARSDROP

Microlanders,” AIAA/USU Smallsat Conf SSC15-XI-3,

DOI 10.13140/RG.2.1.3599.1127.

[2] Klesh A. T. & Halatek L. (2014), International

Astronautical Congress, IAC1-14.B4.8.1.

[3] Klesh A. T. & Krajewski J. A. (2016) “MarCO

– Ready for Launch” CubeSat Developers Workshop

(to be presented 2016/4/21).

[4] Lindensmith CA, Rider S, Bedrossian M, Wal-

lace JK, Serabyn E, et al. (2016) “A Submersible, Off-

Axis Holographic Microscope for Detection of Micro-

bial Motility and Morphology in Aqueous and Icy En-

vironments.” PLoS ONE 11(1): e0147700. doi:

10.1371/journal.pone.0147700

[5] Bekker D. L., Pingree P. J., et al. (2011) “The

COVE Payload – A Reconfigurable FPGA-Based Pro-

cessor for CubeSats”, AIAA/USU Smallsat Conf

SSC11-I-2.

[6] Estlin T. A. et al. (2012) “AEGIS Automated

Targeting for the MER Opportunity Rover,” ACM

Transactions on Intelligent Systems and Technology 3,

Issue 3, Article #50.





ULTRASONICALLY ASSISTED PENETRATION THROUGH GRANULAR MATERIALS. David G. First-brook1, Kevin Worrall1, Ryan Timoney1, Patrick Harkness1, Francesc Suñol2, and Yang Gao3. 1University of Glas-gow, Glasgow, G12 8QQ, U.K., [email protected]. 2Department of Applied Physics, Universitat Politècnica de Catalunya-BarcelonaTech (UPC), c/ E. Terradas, 5, 08860 Castelldefels (Barcelona), Spain. 3University of Surrey, Guildford, GU2 7XH, U.K.

Introduction: Mars is an excellent candidate loca-

tion for historical extraterrestrial life, due to significant evidence pointing to a much wetter and warmer past. The current picture of Mars is very different however, with much of the surface stripped bare due to a reduced atmosphere and high levels of solar radiation. Even the highly radio-resistant organisms D. Radiodurans and Tardigrades are unable to survive over evolutionary time scales at these radiation levels [1]. Nonetheless, the radiation level quickly decreases with depth, reach-ing Earth’s surface levels at roughly 3 m [1].

Gaining access to this depth is a complex task, as the lower gravity on Mars results in a lower available overhead weight. This weight-on-bit (WOB) is an es-sential aspect of drilling and penetrating, with lower values often leading to sub-optimal operation. Power consumption is also a concern, especially for small landers that do not have the volume or mass budget for bulky batteries or solar panels.

A device that can reduce the WOB requirement may allow a probe to travel further under its own (or applied) weight in a reduced gravitational field. Addi-tionally if the device is able to reduce the power re-quired for penetration, this will not only benefit the power-budget of a lander or rover, but could also have a positive knock-on effect on the total mass budget.

This poster will demonstrate that maximum pene-tration force requirements can be reduced by over an order of magnitude compared to non-ultrasonic pene-tration, depending on the amount of ultrasonic vibra-tion used (Fig. 1) [2]. In addition, by utilizing an opti-mum level of vibration, the total power consumption can be reduced (Fig. 2) by up to 28% [3].

New results arising from experiments at the Large

Diameter Centrifuge at ESA’s ESTEC facility (Fig. 3) will also be presented. Here, penetration tests up to 10g were conducted using a variety of ultrasonic ampli-tudes, measuring penetration force as well as consumed power. The results indicate that the benefits of ultra-sonic vibration describe above are likely to be particu-larly applicable in low gravity environments.

References: [1] Hassler D. M. et al. (2014) Science 343 (6169),

1244797. [2] Firstbrook D. G. et al. (2014) AIAA SPACE 2014 Conf. Proc. [3] Firstbrook D. G. et al. (2015) AIAA SPACE 2015 Conf. Proc.

Fig. 1

Fig. 2

Fig. 3



INVESTIGATING TITAN’S GEOLOGY AS MEANS FOR THE PREPARATION OF FUTURE MISSIONS A. Solomonidou1,2, A. Coustenis2, R.M.C. Lopes1, S. Rodriguez3, M.J. Malaska1, P. Drossart2, B. Schmitt4, S.

Philippe4, C. Matsoukas5. 1Jet Propulsion Laboratory, California Institute of Technology, California, USA, 2LESIA - Observatoire de Paris, CNRS, UPMC Univ. Paris 06, Univ. Paris-Diderot France, 3Laboratoire AIM, Université Paris Diderot, Paris 7/CNRS/CEA-Saclay, France, 4Institut de Planétologie et d’Astrophysique de Grenoble, 5Department of Physics, University of Athens, Greece.

Introduction: Saturn’s moon Titan exhibits a dy-

namically-active and nitrogen-rich atmosphere in addi-tion to a complex landscape indicating a rich geologi-cal history. The Cassini spacecraft has been observing Titan during 117 flybys to date obtaining a large da-taset relevant –among other- to Titan’s geological his-tory. Based on the Cassini-Huygens findings our knowledge is now enriched and more targeted concern-ing the future Titan investigations. New landing mis-sions with in situ capabilities are required for atmos-pheric and surface sampling and high-resolution imag-ing. One crucial parameter for a future lander mission will be the selection of an appropriate landing site. In selecting future landing sites one should take into ac-count the science objectives of the mission, the physi-cal properties, the nature of the surface and the engi-neering constraints. Titan’s equatorial zone consists of morphologically and compositionally complex features as shown by the analysis of the Radar Mapper (RADAR) and the Visual and infrared Mapping Spec-trometer (VIMS) data [e.g. 1;2]. These, and a number of more terrains are considered potential landing sites for future in situ missions due to their geological inter-est that could reveal the connection of the surface with the possibly active interior and the atmosphere.

Potential sites of interest: Sotra Patera and Tui Regio, have been proposed as potential sources of me-thane supplied into the atmosphere and/or as areas where photolysis products such as tholin material ac-cumulate, showing a profound connection between the surface and the atmosphere [3;4;5]. In [5] we reported both areas as ‘changing with time’ in terms of surface albedo areas. Both Sotra and Tui are of high interest for future exploration. Other interesting geological units are the geological types called ‘plains’ [6,7]. [6] recently studied several types of ‘plains’ and identified their variable properties. The plains are suggested to be sedimentary in origin, resulting from aeolian or fluvial deposition or accumulation of photolysis products cre-ated in the upper atmosphere. Additional units that require extensive investigation are the dunes, the laby-rinth and the hummocky terrains [7,8]. The study of their organic nature and their link to aeolian/fluvial and or internal processes will significantly contribute to the identification of the nature of the surface of Titan.

Methods: In our study, we fuse Titan data from several Cassini instruments (radar SAR, radiometry,

altimetry, IR spectroscopy, imaging) to determine which areas of the surface present the most scientifi-cally rich targets for future missions in terms of physi-cal and chemical properties. We use specific tools for the treatment of the various data such as a radiative transfer code and despeckle filters [4;5;6;9]. Mission concepts that have been studied include orbiters, bal-loons, landers and even aircraft. Some of these con-cepts include surface sampling. In addition, while stud-ies of geological terrains on a broad scale and liquid-filled lakes and seas have and are still being done, a focused approach to determine promising sampling locations on the basis of combined data sets and acces-sibility has not.

Characterization of terrain types: The geological objectives request landing site measurements of the in situ geological context, chemical composition by sev-eral types of spectroscopy, mineralogy provided by infrared data and petrological properties such as poros-ity, grain size, permeability and more. The nature (composition, texture, homogeneity) of the landing site is of crucial importance especially for the safety of the landing and the sampling of surface material.

Engineering considerations: Among the most im-portant engineering constrains are the precision of landing, the landing ellipse/warning track, the surface hardness, the uniformity, the inclination and the at-mospheric conditions. In situ compositional analysis will require the ability to sample a wide range of pos-sible materials, from loose organic sand grains to fro-zen aqueous cryolavas.

Conclusion: After Cassini-Huygens, future explo-ration could be achieved with measurements taken from close range, at these scientifically interesting are-as, with a montgolfière that could explore the surface in a close-up range and also, perhaps, crash onto the surface and get measurements; and/or a lander that will descend in an equatorial or polar region and take measurements of solid ground or a liquid.

References: [1]Lopes et al. 2010, JGR 118, 416-435. [2]Solomonidou et al. 2013, PSS 77, 104-117. [3]Lopes et al. 2013 JGR 118, 416-435. [4]Solomonidou et al. 2014, JGR 119, 1729-1747 [5]Solomonidou et al. 2016, Icarus 270, 85-99. [6]Lopes et al. 2016, Icarus 270, 162-182. [7]Malaska et al. 2016, Icarus 270, 130-161. [8]Radebaugh et al. 2008, Icarus 194, 690-703. [9]Bratsolis et al. 2012, PSS 61, 108-113.



A MULTI-PLANET, MULTI-SPACECRAFT FLAGSHIP CLASS MISSION CONCEPT TO EXPLORE A

GAS GIANT AND AN ICE GIANT PLANET. A. Mudek1†, and S. J. Saikia2†, 1Graduate Student, amudek@pur-

due.edu, 2Visiting Assistant Professor, [email protected], †School of Aeronautics and Astronautics, Purdue Univer-

sity, 701 W. Stadium Ave., West Lafayette, IN, 47907-2045.

Introduction: In situ probe missions to the outer

planets are designed to answer questions on three pri-

mary themes [1]:

• To constrain models of solar system formation and

the origin and evolution of atmospheres

• To provide a basis for comparative studies of the

gas and ice giants

• To provide a valuable link to extrasolar planetary

systems

The gas and ice giants offer a unique laboratory for

studying the chemistries and dynamics of atmospheres,

and interiors of all the planets, including Earth. The

well-mixed atmospheres and interiors of the giant plan-

ets contains pristine material from the epoch of Solar

System formation that might provide clues to the local

chemical and physical conditions existing at the time

and location at which each planet formed. Uranus, to-

gether with Neptune in our Solar System, belongs to a

class of planets called ice giants. “Ice” in this context

refers to volatile materials that are solid at the low tem-

peratures of the outer Solar System (e.g., water, ammo-

nia, and methane). Therefore, exploration of Saturn us-

ing an entry probe, and that of Uranus and Neptune is

very important to the planetary science community. A

Saturn entry probe is one of the highest priority New

Frontiers mission, highlighted in the Decadal Survey

[1]. At the same time, Jet Propulsion Laboratory, under

the direction of the Planetary Science Division of

NASA’s Science Mission Directorate is study of mis-

sion options, including science and technology options,

for exploring the Ice Giant planets [2].

Spacecraft and Mission Design: The mission con-

cept is based on a unique trajectory discovered at Purdue

University which has a launch year of 2028. The trajec-

tory is an Earth-Earth-Saturn-Uranus opportunity with

one deep-space and the other powered flyby maneuver.

NASA’s Space Launch System (SLS) is used to launch

a collection of high capable spacecraft of mass about

16.5 t in LEO, deliver about 9 t at Saturn, and about 4 t

at Uranus. The time-of-flight from Earth to Saturn is just

5.3 years, and the time-of-flight from Earth to Uranus is

10.5 years.

Due to higher launch capability of the SLS a signif-

icantly large can be delivered, thereby expanding the

types of missions that can be flown. The unique trajec-

tory options such as the Saturn-Uranus one is a game-

changing opportunity. Due to increased delivered

masses, relatively shorter time-of-flight at two planets,

significantly more science can be done in the same mis-

sion (using the same launch vehicle) thereby enabling

new mission concepts.

Figure 1. 2028 Saturn-Uranus mission opportunity

that can deliver a probe at Saturn, and an orbiter and a

probe at Uranus [3].

In this paper, we will present results from an end-to-

end mission concept study (similar to the one done by

NASA JPL’s Team-X) using concurrent engineering

that will include [2]:

Science traceability and definition of measure-

ments, and

Specification of model payload

Detailed orbiter and entry spacecraft design

Define the science/mission concept of operations

Development of mission cost, risk, and program-

matics

Perform trade studies among scientific value, cost,

and risk

Classification of alternative mission architectures,

and

Identification of enabling and enhancing technolo-

gies

The Science Traceability Matrix will be generated

along with JPL’s Ice Giant Study, which will be based

on not only the prioritize science objectives based upon

recommendations from the 2013-2022 Decadal Survey,

but also account for recent information and current state

of the science.

The goal of this study is to aid in the NASA JPL’s

Ice Giant Study as well as to provide information for the



deliberative process for the next, 2023–2032 Planetary

Science Decadal Survey.

References: [1] Space Study Board (2011), Vision

and Voyages for Planetary Science in the Decade 2013-

2022, [2] Green, J. L. (2015), “Planetary Science Divi-

sion Status Report,” OPAG, Aug. 24 Presentation. [3]

K. Hughes et al., NASA OPAG Meeting, Aug. 24.



A Summary Of The SOAREX-8 and SOAREX-9 Sub-orbital Test Flights M. S. Murbach,1 P. Papadopoulos,2 A. Guarneros-Luna,1 R. Alena,1 D. Atkinson, 3 J. Wheless, 1 F.A.Tanner, 3 R.Morrison,2 T. Shu,2 G. Pearhill,3 A. Tabrizi,2 R. Rivas,1 1NASA Ames Research Center, Moffett Field, CA 94035, 2San Jose State University, Aeronautical Engineering Department, One Washington Square, San Jose, CA, 95192, 3University of Idaho, Department of Electrical and Computer Engineering, 875 Perimeter Dr., Moscow, ID, 83844

Abstract: Two sub-orbital missions, SOAREX-8 and SOAREX-9 (Sub-Orbital Aerodynamic Re-entry Ex-periments) were conducted on July 7, 2015 and on March 7, 2016 from the NASA Wallops Flight Facility test range. Each achieved an apogee of 332 and 160 km, respectively. The first one jettisoned a small-sat payload before apogee, carrying a full-scale Exo-Brake de-orbit device, as well as set of precursor telemetry and data acquisition devices (command/control teleme-try; 1Mbs data rate solution; 10 Mbs data rate solution) for use on future test flights. The SOAREX-9 experi-ment further tested the data acquisition systems, as well as the Generation 2 Wireless Sensor Modules (WSM) with notable success. The payload was also retrieved after the flight from the ocean environment with additional data recovered. The Exo-Brake fabrica-tion, testing and validation are discussed – and also as a viable approach for de-orbiting payloads of varying sizes. These experiments help to set the stage for up-coming orbital experiments that will be de-orbited from the International Space Station. The application of these data-acquisition, tracking, and wireless sensor network techniques - to future ascent and re-entry ex-periments, will be described.



ASSESSMENT OF VENUS AERIAL PLATFORMS FOR IN-SITU MEASUREMENTS AND ANALYSIS OF OPTIMAL ALTITUDE FOR SUPERPRESSURE BALLOONS. R. J. Rolley1†, Y. Lu2‡, S. J. Fulton, Jr.3‡, E. Shibata4‡, J. A. Cutts5, M. de Jong6, S. J. Saikia7‡, 1Undergraduate Student, [email protected], 2Ph.D. Candi-date, [email protected] 3Undergraduate Student, [email protected], 4Graduate Student, [email protected], 5NASA-Caltech Jet Propulsion Laboratory, 4800 Oak Grove Drive, Pasadena, CA, 91109, [email protected], 6Thin Red Line Aerospace, 208-6333 Unsworth Rd, Chilliwack BC V2R 5M3, Canada, [email protected], 7Visiting Assistant Professor, [email protected], †School of Mechanical Engineering, Purdue University, 585 Purdue Mall, West Lafayette, IN, 47907, ‡School of Aeronautics and Astronautics, Purdue University, 701 W. Stadium Ave., West Lafayette, IN, 47907.

Exploring Earth’s Twin—Scientific Impetus for

Venus Missions: Exploring the dichotomy between Venus’ twin-like similarities and stark contrasts to Earth presents a unique opportunity to answer some of the most fundamental questions posed by science. Ve-nus has a size and bulk density similar to that of Earth, suggesting that the two planets formed in much the same way in the early Solar System. However, subse-quently the two planets followed different evolutionary paths [1], as present day Venus is completely inhospi-table—harboring surface pressures higher than 96 bar and temperatures reaching 740 K [2]. Investigation of Venus can help to understand the evolution of terres-trial planets and identify the conditions necessary to maintain habitable environments on planetary bodies. Additionally, Venus science missions can provide an-swers to critical yet unanswered questions about Earth—why conditions evolved to be habitable, how long they will remain habitable, and how our climate system operates and changes over time [1]. As outlined by the National Aeronautics and Space Administration (NASA) Venus Exploration Analysis Group (VEXAG), based on the findings of the Planetary Sci-ence Decadal Survey: Visions and Voyages for Plane-tary Science in the Decade 2013-2022, future missions to explore Venus should seek to address the following high priority goals: [1] • Understand atmospheric formation, evolution, and

climate history on Venus • Determine the evolution of the surface and interior

of Venus • Understand the nature of interior-surface-

atmosphere interactions over time, including whether liquid water was ever present Compelling Nature of Aerial Platforms: Orbiter

missions such as Pioneer Venus, Magellan, Venus Ex-press, and Akatsuki have collected data to create de-tailed global surface maps, measure surface emissivity, and determine some atmospheric constituents [3]. However, the science return of future missions to study Venus can be dramatically increased by the use of in-situ measurement platforms. While current technology

exists to fabricate short-duration landers, significant technology investment is required to enable the use of a long-duration surface lander [1]. In contrast, aerial platforms hold significant potential for use in Venus missions due to their ability to conduct in-situ meas-urements of temperature, pressure, wind speed, and chemical composition for extended durations—ranging from tens of days to a year—in a more benign envi-ronment than the Venus surface. These platforms should therefore be considered to meet the goals of VEXAG and the Planetary Science Decadal Survey [2].

Looking to the Future of Venus Exploration: To date, the VeGa balloons developed and launched by the Soviet Union in 1984 remain the only long duration aerial exploration platforms flown at Venus [4]. As NASA, the European Space Agency (ESA), Indian Space Research Organisation (ISRO) and other space agencies discuss the options of returning to Venus with an aerial platform, an analysis of several prominent platforms shall be conducted. In this study there will be a review of the technical specifications, science capabilities, and current state of development of these platforms as well as a trade study aimed at determining the optimal altitude to fly a superpressure balloon on Venus. The proposed work will present the results of these analyses to provide a reasonable comparison of the selected platforms, so that future mission designers can more quickly access the most feasible craft for their desired mission.

Comparative Aerial Platform Study: Numerous aerial platforms have been proposed for Venus explo-ration missions, and one platform—VeGa—has al-ready been flown. A list of these key platforms is in-cluded below. • VeGa superpressure balloons—flown by Soviet

Union • Venus Aerostatic Lift Observatories for in-situ Re-

search (VALOR) superpressure balloon—proposed by NASA Jet Propulsion Laboratory (JPL)

• European Venus Explorer (EVE) superpressure balloon—proposed by ESA



• Altitude-adjustable balloon based on Ultra High Performance Vessel (UHPV) technology—proposed by Thin Red Line Aerospace

• Metallic bellows balloon—proposed by NASA JPL • Venus Atmospheric Maneuverable Platform

(VAMP) semi-buoyant powered glider—proposed by Northrop Grumman

• Venus airplane concept—proposed by NASA Glenn

These platforms are pictured in Figure 1.

Figure 1: Proposed and flown aerial platforms for Venus include altitude-adjustable balloons, static superpressure balloons, metallic bellows balloons, gliders, and planes (Ve-hicles not to scale) [5][6][7][8][9].

A detailed comparison of the overall dimensions, system and payload mass, operating range—altitude, latitude and longitude reachability—lifetime, potential science measurements, and overall technology readi-ness level will be conducted to produce figures of mer-it for each platform. The figures of merit will provide the basis for comparisons of risk, complexity, science capability, and system design to inform what science objectives are achievable with current technologies, provide insight into which platform(s) should be used for near-term exploration missions, and aid in the next 2023-2032 Planetary Science Decadal Survey.

In addition to conducting an assessment of a broad suite of aerial platforms, we also plan to complete a trade study focusing on the operation of a specific type of aerial platform: superpressure balloons.

Balloon Altitude Trade Study: The VeGa super-pressure balloons launched by the Soviet Union flew at an altitude of 54 km. Similarly, the VALOR and EVE superpressure balloons are being designed for nominal float altitudes of 55 km. A trade study of superpressure balloon flight at various altitudes in the Venus atmos-phere will help to determine if there is an optimal alti-tude or altitude range to fly balloons on Venus other than the 54–55 km range used by past and proposed platforms. Considerations in this trade study will in-clude upwelling and downwelling solar and thermal radiation balance, temperature, pressure, and density conditions, atmospheric and cloud constituents, and the logic of the 54–55 km range selection of proposed plat-

forms. To conduct this study, parametric models of spherical superpressure balloons will be developed and used to assess trade-offs between balloon volume and mass, payload mass, buoyancy, and altitude, to deter-mine optimal flight altitude to maximize science return given current balloon technology readiness, cost, and risk.

In addition to the comparative analysis of super-pressure balloons, we will examine the trades for those concepts—such as metallic bellows balloons—which can descend much deeper in the atmosphere including the availability of solar power, electronics performance and buoyancy considerations.

References: [1] Herrick R. et al. (2014) Goals, Ob-jectives, and Investigations for Venus Exploration (VEXAG). [2] Coustenis A. et al. (2010) Journal of Aerospace Engineering, 225. [3] (2014) Venus Explo-ration Themes (VEXAG). [4] Stetson D. et al. (2014) Roadmap for Venus exploration (VEXAG). [5] Baines K. et al. (2012) Exploring Venus with Balloons – Sci-ence Objectives and Recent Technical Advances (IPPW 9). [6] Lee G. et al. (2014) Exploring Venus’ Atmosphere with a Semi-buoyant Air Vehicle (IPPW 11). [7] de Jong M. (2015) Venus Altitude Cycling Bal-loon: An Introduction to UHPV (Venus Lab and Tech-nology Workshop – VEXAG). [8] Landis G. A. (2002) Solar Airplane Concept Developed for Venus Explora-tion. [9] Balint T. et al. Venus Aerial Mobility Con-cepts: In the Clouds & Near the Surface.



ATHENEA - A mission proposal to a Main Belt Comet. C. M. Mockel1 , J. Berzosa Molina1, P. Vizcaíno Rubio1 and M. Coppola2, 1 Delft University of Technology, Department of Astrodynamics and Space Missions (Kluyverweg 1, 2629HS Delft, Netherlands, [email protected] ), 2 Delft University of Technology, Department of Control and Simu-lation.

Introduction: Asteroids that have not undergone

thermal processing within their lifetime are contempo-rary witnesses of the conditions found in the early So-lar System, and thus hold invaluable information.

This article proposes ATHENEA (Astro-biology, Tides and Hazards Explorer of a New Elementary As-teroid), a robotic M-class mission with the primary objective of analyzing an active asteroid (bodies with cometary characteristics but with orbital trajectories typical of asteroids).

ATHENEA's target asteroid is P/2013-R3, deemed highly favorable for its recent fragmentation into 10 smaller pieces (forming three clusters) [1]. This allows the measurement of interior properties (virtually unaf-fected by space weathering) without the need for pene-tration tools.

ATHENEA consists of a mothership with 8 cu-besats. The mothership transports the cubesats within reach of their target destination. Six cubesats are to be used for P/2013-R3 and two are reserved for use on a Near Earth Asteroids (this secondary mission target would be reached via a fly-by during ATHENEA's interplanetary travel). Cubesats are favored because it is then possible to distribute measurement instruments and potentially split up to observe different sub-bodies of P/2013-R3. Mission critical and resource intensive instruments remain hosted on the mothership, which also acts as a relay station for communication back to Earth.

The instruments on board of the mothership in-clude: visual camera, thermal infra-red camera, spec-trometer, dust sensor, deep radar, and topographer. The six cubesats destined for P/2013-R3 include (distribut-ed among each other): magnetometer, dust sensor, VEGA, spectrometer, plasma sensor, infra-red micro-scope, visual camera, beacon tracking. The two cu-besats destined for the Near Earth Asteroid are still to be defined; they could be used for secondary objectives by external (research/commercial) partners.

The mothership can gather data on one element of the ≈ pieces from P/2013-R3 whereas the cubesats (with the exception of one which includes the topogra-pher) would focus on a piece from another cluster from the break-up. The launch date would be around 2028, and the interplanetary travel time is expected to last approximately 5.5 years. The mission has a designed lifetime of 3 years, which is the expected usage life-time of the cubesats.

References: [1] D. Jewitt, et al.(2014), “Disinte-

grating asteroid p/2013 r3,” The Astrophysical Journal

Letters784, L8



mailto:[email protected]?subject=ATHENEA%20-%20

Biosignature Explorer for Europa (BEE) Probe –Directly Searching for Life Evidence on Europa. Michael J.

Amato1, P. Spidaliere1, P. Mahaffy1, C. Schiff1, O. Hsu1, T. Hurford1, M. Benha1, W. Brinckerhoff1, J. Garvin1, J.

Downing1, T. Errigo1, D. Glavin1, M. Sarantos2, R. Lorenz3, T. Hoehler4, et al. 1NASA Goddard Space Flight Center

(GSFC), Greenbelt MD. 2 University of Maryland Baltimore County/NASA Goddard Space Flight Center, 3 JHU

Applied Physics Laboratory, 4 NASA Ames Research Center,

Introduction:

Evidence for Europa’s substantial sub crustal water

ocean leaves us with the curcial question of if the

ocean is habitable and if it harbors life.

The tantalizing possibility of hydrothermal activity

below its ice crust shell has placed Europa as one of

the highest priority targets in the search for habitable

environmentsand potentially extant life [NRC, 2011].

The compositional analysis of potential plumes is a

high-value objective. Sporadic eruptive plumes at Eu-

ropa have been observed [Roth et al., 2013], and analy-

sis shows plume activity is likely at lower intensities

The plumes represent an opportunity to uniquely-

probe the chemistry of the subsurface ocean and assess

Europa’s potential to sustain life by looking for bio

signatures. Freshly-ejected ocean material would repre-

sent a relatively unaltered sample of the subsurface

chemistry, as compared to sputtered surface materials

exposed to high energy radiation that will destroy or-

ganic compounds. In addition to the detection of more

pristine organic molecules, the composition of the oth-

er constututnes such as salts in the ocean has important

implications for the source of ocean materials [e.g.,

Brown and Hand, 2013].

The potential changing intensity of the plumes at

Europa calls for a versatile measurements strategy that

can accommodate a wide range of geographical loca-

tions and outgassing intensities. While the NASA Eu-

ropa mission spacecraft is well equipped for Europa

studies, which include plume detection and characteri-

zation, more is needed to fully analyze Europa plumes.

The recently selected instrument suite likely lacks the

ability to achieve the highly desirable and difficult de-

tection of direct biosignatures or life evidence. In addi-

tion, the Clipper mission may not be able to easily

modify its orbital trajectory or altitude to fly through

plumes or remotely study plume events that may be

short lived or highly diffuse. Thus, a smaller Probe

equipped with the focused set of instruments and navi-

gation capabilities is better suited to target detection of

bio signatures and can target denser plumes and lower

altitudes that might be otherwise inaccessible to the

Europa Clipper spacecraft.

The BEE plume probe:

A small plume probe would have more flexibility to

perform the critical science investigations in support of

the goal of exploring Europa plumes for evidence of

past or current life on Europa. Objectives to meet this

goal could include: 1) Characterizing the building

blocks of life within the plumes, 2) investigating plume

source regions to assess them as a biological niche en-

vironment and 3) further assess plume source regions

for future landed missions.

Our team has designed a Biosignature Explorer for

Europa (BEE) Probe concept to ‘taste’ the ocean by in

situ analysis of plume sample for biosignatures, which

provide the most science and most programmatically

robust way to determine if the Europa ocean harbors

life evidence. By flying directly through a plume, the

BEE is able to sample freshly released ocean water and

search for evidence of extant life. BEE does this by

using newly matured collection approaches and mass

spectrometer designs based leveraging heritage ap-

proaches. The search for direct in situ molecular bi-

osignatures is the clearest path toward to definitive

identification of signatures of life as we know it at Eu-

ropa.

The BEE will search for molecular signatures of

life by capturing material soon after it is released from

a subsurface reservoir, and conduct chemical analysis

with experiments optimized for detection of molecular

biosignatures. Plume fly-through with sampling and

molecular analysis is akin to “landerless landing” in an

ocean (or fresh ocean deposit) in the quest for Europa

ocean biosignatures

The BEE fly-through approach may offer many

advantages over static landings. BEE plume sampling

collects intact biomarker material before their inevita-

ble destruction by radiation after release from the

ocean. BEE offers better access to freash material by

targetating plumes or active targets (2-10 km fly

through corridor vs. lander Braking-Descent-Landing

(BDL) uncertainty of at least 75-100km or worse).

BEE is less sensitive to radiation damage and effects

on sensitive instrumentation. BEE has no landing loads

on sensitive bio-molecule sensors. BEE’s broad re-

gional context can be established with simple imaging

systems.

BEE can be more agile and responsive to activity

on Europa when we arrive with the Europa mission,



allowing easier targeting to the “the action”. BEE has

no BDL lander systems and cost, uses less resources,

has less complexity, has fewer mission constraints and

no landing risks that will drive additional costs. BEE

has an advanced NASA Goddard Space Flight Center

led mass spectrometer with a large area collection ap-

proach that leverages past designs while enabling new

direct biosignature science. The sensitive mass spec-

trometer is combined with other separation approaches

for definitive identification of biosignatures. It also

carries the BEE UV plume targeting camera as well as

visible and IR cameras to image the active region with

better resolution than the Europa mother ships instru-

ments.

The BEE team has refined its cadence of prerelease

survey, probe release, refined targeting, sample collec-

tion, analysis and transmit operations. BEE is released

from the mother ship many hours before closest ap-

proach to Europa. Initial targeting is done using plume

data from the Europa mission instrument suite. Refined

targeting on a plume is aided by BEEs unique targeting

camera that senses emissions night or day. The BEE

probe flies thru at very low km altitude and collects

material. The probe images the surface in visible and

infrared at process-diagnostic scales. After exiting the

intense radiation environment, it uses proven mass

spectrometer technology to analyze the material for

biosignatures. The BEE then transmits the data back to

the Europa mission mother spacecraft. The BEE can be

released during a large number of the baselined flybys.

The BEE uses its refined targeting and propulsion sys-

tem to enable targeted access a majority of Europa’s

surface.

The small BEE probe attaches to the Clipper mis-

sion, currently on its NADIR viewing side. The BEE

team has worked with the Europa mission project engi-

neering team to work out initial mechanical, load, elec-

trical, communications and operations details that have

low cost and resources impact on the main mission.

BEE employs a modular mechanical design and a care-

fully designed internal radiation protection vault to

protect sensitive electronics from radiation effects and

is under 250 Kg.

The BEE probe ACS and propulsion system are de-

signed to enable targeting and post sample collection

maneuvers. BEE is three axis stabilized and utilizes

acceleration measurement systems, star sensors and

other packages. The propulsions system is a hybrid

system combining a basic bi-propellant design with a

simple hydrogen driven three axis cold gas system that

is compatible for use just before and during sample

collection.

The BEE’s avionics utilize a selectively redundant

design that leverages multiple NASA Goddard Space

Flight Center (GSFC) systems designed for far from

earth applications. The design has high performance

processor and memory units combined with a number

of interfaces cards. These interfaces include power and

data interfaces, propulsion and actuator interfaces, re-

fined targeting sensor interfaces and redundant X-band

transponder. The BEE’s power system is currently a

heritage primary battery designed with multiple battery

cell strings.

The BEEs thermal system utilizes blanketing and

four thermal zones as well as survival heaters. The bat-

tery system is isolated from the electronics vault to

minimize heating requirements but the propulsion tanks

are thermally linked to the vault. A separate thermal

interface will be linked to the mother spacecraft ther-

mal system during cruise.

The BEE probe is a feasible way to achieve criti-

cal biosignature and life evidence goals with less risk

and lower cost than other options. The BEE team has

shown this approach to be feasible approach to be con-

sidered.

Figure 1 – The BEE probe design and its current loca-

tion on the Europa mission spacecraft.

Acknowledgments: The author would like to

acknowledge NASA Headquarters Planetary Science

Division PSD for follow on study support and the

many team members at NASA GSFC and other institu-

tions involved in the BEE probe work.

References - [1] NRC, 2011, [2] Roth et al., 2013 [3]

Brown and Hand, 2013



Concept of Lunar Ballistic Robot Exploration (BARE)

Balaji Soundararajan

Technical University of Berlin, Marchstr 12-14, 10587, Berlin, Germany, email: [email protected]

Introduction: The motivation of this paper is to

present about the possible expansion of the lunar ex-

ploration capabilities with Human presence on the

Moon using a Ballistic type Robot named BARE.

Though this is a futuristic topic, it is important to con-

sider ways of exploration of Lunar surface. With major

Space organisations around the world wanting to go

back to the Moon in 2020s, it is necessary to find in-

novative methods of exploration.

Mission Principle. The Ballistic type Robotic ex-

ploration will make use of the very less gravitation and

absence of atmosphere on the moon. Once launched

from a point by an astronaut, the BARE robot will fol-

low a ballistic parabolic trajectory and heads to farther

distances. It is based on the fire and forget type launch.

It is initially launched from a cannon type Ballsitic

Robot launcher. With the help of small solid booster at

the back, fine attitude adjustments are made with re-

spect to velocity and altitude. Once it reaches a par-

ticular altitude based on initial velocity, the BARE

heads down in a powerless flight. It hard lands at a

specific location. The BARE is intended to be a small

cylinder shaped robot. Once it lands at the intended

area (with 90% accuracy), the BARE opens up the

front door and releases small ball shaped robots that

are tethered to this parent cylinder shaped robot. These

small ball shaped robots are 3 to 10 in number and

carries out Insitu experimentation of regolith. The par-

ent robot also opens up two solar panels for power

generation and a radio antenna for communication.

Figure 1 below shows the operation of a BARE.

Figure 1. Operation of BARE

System design. The BARE is a small cylindrical

type robot of 80 cm in length and 45cm in diameter.

Two solar panel opens up after hard landing using

springs. A radio antenna also helps in communication

with station at Lagrangian points or satellites in lunar

orbits. One reaction wheel is used for Roll control of

the BARE robot to make sure it lands using those

springs. Ball shaped robots carry out real time surface

based measurements and experimentation. Each ball

robot has a specific task to do. They can be customized

and are powered using these tethers. The entire BARE

is assembled by the astronaut from appropriate items

carried from Earth. The HTPB based solid propellent

with a small booster at the back is intented to have a

burn time of 5 sec initially to impart proper required

velocity from the launcher. Each BARE robot has an

identification number and many such robots are

launched at various locations. Information from each

BARE is collected and recorded.

Possible Application. With an initial velocity of

2.00Km/sec, the BARE reaches a maximum altitude of

310Km and travels as far as 2100 Km. This means we

will have the capability to explore farther distances

from the lunar outpost. The real time experiments will

be carried out on long term basis and any vital or inter-

esting discoveries by these robots will be of great help

for next manned missions, and planning of next out-

post can be chosen. As a further advancement in this

technology, we can use these in Martian or other planet

environment too, to launch autonomously. These ro-

bots can be sent to far side of the moon without solar

panel, but with long term battery life.

Conclusion. With robots like this, it is possible to

explore the Lunar surface at a bigger scale, given the

opportunity to explore. With the astronauts available to

operate, this kind of robot will open the gate for a new

means of exploration. We will have the opportunity to

have these robots almost all over the surface of the

planet. This will surely prove to be Exploration friend-

ly vehicle.



Cubesat Application for Planetary Entry (CAPE) Missions: Micro-Return Capsule (MIRCA). J. Esper, C. A.

Nixon, M. Amato, J. B. Garvin, NASA Goddard Space Flight Center, Greenbelt MD 20771.

Introduction: So far, no microprobe (less than 10

kg) has entered another planetary atmosphere and suc-

cessfully relayed data back to Earth. Although the

Deep Space 2 Mars microprobes did reach their desti-

nation (total mass about 6.5 kg each), unfortunately

they were lost due to a combination of delivery system

failures and other unknown factors. This paper de-

scribes a planetary entry probe based on the widely

popular Cubesat-class spacecraft specification (Cubesat

Application for Planetary Entry Missions, or CAPE

probes) [1], [2]. CAPE not only opens the door to new

planetary mission capabilities, it also offers relatively

low-cost opportunities especially suitable to university

participation.

CAPE: The CAPE concept describes a high-

performing Cubesat system which includes a propul-

sion module and miniaturized technologies capable of

surviving atmospheric entry heating, while reliably

transmitting scientific and engineering data. In broad

terms, CAPE consists of two main functional compo-

nents: the “service module” (SM), and “CAPE’s entry

probe” (CEP). The SM contains the subsystems neces-

sary to support vehicle targeting (propulsion, ACS,

computer, power) and the communications capability

to relay data from the CEP probe to an orbiting “moth-

er-ship”. The CEP itself carries the scientific instru-

mentation capable of measuring atmospheric properties

(such as pressure, temperature, composition), and em-

bedded engineering sensors for Entry, Descent, and

Landing (EDL) technology monitoring and assessment.

Figure 1 shows the complete vehicle system.

Serv

ice

Mo

du

le

Deployed Array

Micro-thruster

UHF or S-Band Comm.

2U

Cu

beS

at

~1U Microprobe

Pla

net

ary

Entr

y P

rob

e

Figure 1: CAPE in its deployed configuration, and

stowed inside deployment system

MIRCA: In order to reduce CAPE’s implementa-

tion risks, the Micro Return Capsule (MIRCA) CEP re-

entry demonstrator is currently being prototyped at the

NASA Goddard Space Flight Center (GSFC). Figure 2

shows MIRCA inside a wind tunel at the Wallops

Flight Facility (WFF), where laminar flow and aerody-

namic stability were demonstrated at low speeds.

Figure 2: MIRCA in wind tunel test at NASA WFF

(shown at angle of atack ≈ 20°).

Flight Test No. 1: The first flight of MIRCA was

successfully completed on 10 October 2015 as a “pig-

gy-back” payload onboard a NASA stratospheric bal-

loon launched from Ft. Sumner, New Mexico. This

completed verification of its avionics, including the

Inertial Measurement Unit (IMU), single board com-

puter, power conditioning and distribution system,

UHF communications transceiver, on-board thermal

sensor, (ground) telemetry acquisition system, and

flight software, all critical steps in MIRCA’s develop-

ment. Figure 3 shows a view from 30 km altitude, as

well as data from the IMU during gondola release and

parachute depoloyment.

Figure 3: Acceleration and body rate data recreate

vehicle dynamics. Shown is maximum axial (+Z-axis

vertical) deceleration of about 3.3g after parachute

deployment.



In the first flight test, MIRCA remained tethered to

the gondola, and was not released. A second “drop

test” is planned for later in 2016 to verify its aerody-

namic performance at velocities approaching the speed

of sound.

Applications: The CAPE/CEP architecture could

provide an important new mechanism for probing plan-

etary atmospheres, such as those of Mars, Venus and

Titan. If equipped with a deceleration system for con-

trolled descent (e.g. parachute), a number of probes

could be deployed to make simultaneous measurements

of the vertical atmospheric structure, much in the same

way as sounding rockets are used on on the Earth, but

in reverse. This type of approach was previously used

in the Pioneer Venus mission, which deployed three

small and one large instrumented atmospheric probe.

The Huygens lander on Titan, and Galileo Jupiter

probes performed similar function, but provided only a

single vertical sampling. The advantage of CAPE/CEP

therefore is in providing the possibility of multiple,

simultaneous atmospheric measurements.

The probe measurements could include: tempera-

ture, pressure, wind shear (and perhaps directional ve-

locity, if combined with accurate tracking), net flux in

visible and infrared channels, and perhaps some meas-

urements of composition, especially for more abundant

gas species (e.g. H2O, CO2 on Mars and Venus, CH4 on

Titan). These measurements would provide an im-

portant test of dynamical models of the atmosphere.

References: [1] Esper, J., Baumann, J-P., Herdrich,

G.: Cubesat Application for Planetary Entry Missions

(CAPE), Interplanetary Small Satellite Conference,

California Institute of Technology, Pasadena CA, June

2013. [2] Esper, J.: Cubesat Application for Planetary

Entry Missions (CAPE): Micro-Return Capsule

(MIRCA), International Workshop on Instrumentation

for Planetary Missions (IPM-2014), NASA Goddard

Space Flight Center, Greenbelt, Maryland, November

2014.



Exomars Heatshield

Yann Mignot, Aurélien Pisseloup, Jean-Marc Bouilly - AIRBUS DS

[email protected], [email protected], jean-

[email protected]

On March, the 14th, EXOMARS 2016 Spacecraft was launched successfully from Baïkonour by a

Proton Launcher.

After seven months of cruise, Schiaparelli Entry Descent & Landing (EDL) module will separate from

TGO (Trace Gaz Orbiter), and on the 19th of October, while TGO will insert in Mars Orbit, Schiaparelli

will start its Entry in Mars atmosphere. This will be the time for Schiaparelli’s heatshield to protect

the probe from the stringent aerothermal conditions as the speed is reduced in a few minutes from

21000km/h to 240km/h, before the landing platform goes on its own for the final phases of the

descent.

In the presentation, AIRBUS DS will firstly make a focus on typical features of Schiaparelli’s

Frontshield and Backshield, before giving an insight into the development, qualification, and

manufacturing of the Heatshield, up to the very last AIT activities performed during Launch

campaign.







GLIDING THROUGH THE HYDROCARBON LAKES OF TITAN USING A STEERABLE PARACHUTE. Y. Lu1†, M. de Jong2, and S. J. Saikia3†, 1Ph. D. Candidate, [email protected], 2Thin Red Line Aerospace, 208-6333 Unsworth Rd, Chilliwack BC V2R 5M3, Canada, [email protected], 3Visiting Assistant Professor, [email protected], †School of Aeronautics and Astronautics, Purdue University, 701 W. Stadium Ave., West Lafa-yette, IN, 47907-2045.

Introduction: The moon of Saturn—Titan, is the only celestial body other than Earth known to harbor stable surface liquids. It is also the only world in the So-lar System beyond Earth that is shielded from deadly radiation, and harbors a benign, yet extremely cold, sur-face environment with nitrogen-dominated atmosphere [1]. Titan’s atmosphere experiences a wide range of dy-namical and chemical processes, and varies seasonally in temperature and composition. Surface liquids act as a natural collection system for the global organic inven-tory, and noble gases which upon investigation could provide new insights into organic chemical evolution on a global scale and detect ongoing biological processes [1]. Most surface liquids are in the lakes and seas in the northern hemisphere (refer to Fig. 1): Kraken Mare (1200 km), Ligeia Mare (500 km), Punga Mare (380 km). There is only one lake in the southern hemisphere, Ontario Lacus (235 km by 75 km). These bodies of liq-uids are of very high interest for future exploration. Figure 1. (a) Northern seas and lakes (Kraken, Ligeia, and Punga), (b) the only southern lake, foot-shaped Ontario Lacus (Source: NASA/ESA, Cassini Radar Mapper)

Both Titan Mare Explorer (TiME) studied by John-Hopkins University and Titan Saturn System Mission (TSSM) by Jet Propulsion Laboratory (JPL) in the de-cadal survey proposed to land on Kraken (Ligeia) Mare; however, the window of opportunity is very narrow for Direct-to-Earth (DTE) communication in the northern hemisphere (Ligeia and Kraken are only visible from Earth before 2027, whereas Ontario lacus will be visible from 2025-2040) [2], thus Ontario Lacus appears to be an interesting target for lake landers.

Huygens Probe Entry, Descent, and Landing Huygens probe successfully landed on Titan in 2005, via ballistic entry using a rigid aeroshell (60° sphere-cone) followed by a sequence of three non-steerable

parachutes for descent and touchdown. The sequence of non-steerable parachute used in Huygens is shown in Fig. 2 [3]. A pilot chute is first deployed to remove the rear aeroshell, followed by the deployment of Disk-Gap-Band (DGB) main parachute for front shield re-lease and descent phase for science measurements. Due to the limited lifetime (3 hours) of Huygens’ on-board battery, as the descent would last for more than three hours using only the main parachute. a third drogue par-achute is deployed for faster descent until final touch-down.

Figure 2. Sequence of Huygens probe descent control sub-sys-tem: (a-c) pilot parachute is deployed and rear areoshell sepa-rated, (d-f) main parachute is deployed and front shield sepa-rated, (g) drogue parachute is deployed for final descent. [Un-derwood, 2006]

Why do we Need Steerable Parachute: Due to the atmospheric uncertainties on Titan, such as density, pressure, and wind, a rigid aeroshell entry system pre-sents a large error ellipse. In addition, a non-steerable parachute further increases the uncertainties in the error ellipse. For instance, the size of the landing error ellipse is larger than the size of Ontario Lacus, thus an entry vehicle cannot be guaranteed to land on the lake. For future Titan exploration, it is important to land accu-rately at or near scientific important landing sites—On-tario Lacus, shoreline, river, a particular area in Kraken Mare, near a cryovolcano, etc.

In order to improve the landing accuracy, either on hydrocarbon lakes or on surface, guidance (similar to Apollo capsule and Mars Science Laboratory entry ve-hicle) during entry, descent, and landing (EDL) will be required. Since a considerable period during the EDL timeline on Titan is using a parachute, guidance will be required not only for rigid aeroshell but also for the par-

400 km

a b



achutes. Therefore, steerable parachutes will be re-quired for future Titan missions. No previous missions have employed steerable parachutes for probe descent and landing. In the proposed study, we will survey the different designs of steerable parachutes on Earth that can be extended to space exploration.

Steerable Parachute: Steerable parachute (e.g. parafoil in Fig. 3a), was demonstrated by X-38 [4] which successfully steered the craft to the desired land-ing target and reduced the landing speed of 128 m/s to 20 m/s. Such parafoil is also popular among extreme sport, e.g. parachuting. There are several types of steer-able parachutes that have been tested and proven for Earth applications. Some of the steerable parachutes are:

Figure 3. Selected steerable parachute designs: (a) X-38 para-foil, (b) MC-6 parachute system, and (c) Clustered round par-achute. 1. Parafoil (ram-air parachute). Resembling a low as-

pect ratio wing, asymmetric manipulation of trail-ing edge control lines creates asymmetric drag and resulting maneuverability. Example: Joint Preci-sion Airdrop System (JPADS) [Ref: DoD].

2. Modified circular parachutes, such as Airborne System’s MC-6 Parachute System (Fig. 3b), Af-fordable Guided Airdrop System (AGAS) [5]. Such systems are maneuvered via vented canopy and control input.

3. Clustered round parachute (Fig. 3c) [6], separate suspension lines are attached to each parachute, and pulling the lines will create a difference in the force generated by each round parachute.

4. Parawing (inflatable/flexible ‘Rogallo’ wing), is considered for Gemini and Apollo spacecraft [2], is essentially a flexible wing/airfoil.

Lift to drag ratio (or glide ratio) is an important pa-rameter for steerable parachutes, Table 1 summarizes the controllability and typical range of L/D ratio for se-lected steerable parachutes. The table shows that all the different systems have both controllability in out-of-plane and in-plane motion. However, there exist signif-icant differences in the maximum banking rate, and ad-justable range of glide ratio.

Table 1. Controllability and L/D ratio for selected steerable descent system

System Controllability L/D ratio In-plane Out-of-plane Parafoil Downrange Bank 1.8 to 6 MC-6 Downrange Bank 0.8 to 0.9 AGAS Longitudinal Lateral 0 to 0.5

Clustered Downrange Bank 0.3 to 0.9 Parawing Downrange Bank 10 to 20

In the proposed study, different designs of steerable parachutes will be assessed based on the system and payload mass, reliability, packaging, deployment, and controllability. The performance and dynamics of dif-ferent steerable parachutes for Titan environment will also be presented, and the results will be compared for both Titan and Earth environment. Effects of gravity and atmospheric conditions (wind, density, and pres-sure) on the parachute dynamics will be studied.

References: [1] Space Study Board (2011), Vision and Voyages for Planetary Science in the Decade 2013-2022, [2] Lorenz, R. D., and Newman, C. E. (2015) Adv. in Space Research, 56, 190-204. [3] Underwood, J. C., (2006) Huygens DCSS performance Reconstruction, IPPW-3, [4] Lingard, S., and Underwood, J., (2014) En-try, Descent and Landing System, Springer, [5] Del-licker, S., et al., (2003) AIAA 2003-2101, [6] Lee, C. K., and Buckle, J. E. (2004) Journal of Aircraft, 41, 5, 1191-1195.



Gravity Tractor Dynamics under the Effect of Non-Homogenous Gravity Fields of Asteroids. S. P. Shekhar1 ([email protected]) and Y. Ketema1 ([email protected]) , University Of Minneasota, Minneapolis, Minnesota.

Abstract:The study of Near-Earth objects[1,2] is very important as it can pose a greater threat to Earth. Sev-eral theories for Asteroid Deflection have been sug-gested to avert an impact[3,4]. The gravity tractor is one of the method discussed most is used to lessen such risks[5]. Gravity tractor is a spacecraft which works in close proximity of an asteroid uses mutual gravitational force to perturb the asteroids from its state to avoid impact. The method proposed here for deflection of asteroid that make use of spacecraft mov-ing back and forth on a segment of an appropriate Kep-lerian orbit about the asteroid. It is shown that on aver-age the spacecraft describing such a trajectory can ex-ert a significantly large force on the asteroid than a stationary gravity tractor[6-8] thereby reducing the time needed to effect the desired velocity change. The difficulties in designing the orbits for spacecraft are addressed to maintain an equatorial plane by study-ing the nonhomogeneous gravity fields of rotating asteroids. The back and forth motion around the aster-oid allow us have a periodic orbit which eases the tug-ging effect. Furthermore , the current method does not require canted thrusters[6] on the spacecraft, markedly reducing the amount of fuel needed for a given veloci-ty change. In addition, the method allows for the sim-ultaneous use of several spacecraft,further strengthen-ing the overall tugging effect on the asteroid which indeed reduces the mission time. The method we ap-plied in designing the orbits can be used in the applica-tions like Navigation and Communication around as-teroid.

References: [1] Chapman, C. R., and Morrison, D., “Impact on the Earth by Asteroids and Comets: Assessing the Hazard,” Nature (London), Vol. 367, No. 6458, 1994, pp. 33-39. [2] Binzel, R. P., “The Torino Impact Hazard Scale,” Planetary and Space Science, Vol. 48, No. 4, 2000, pp. 297-303. [3] Izzo, D., Olympio, J., and Yam, C., “Asteroid De-flection Theory: Fundamentals of Orbital Mechanics and Optimal Control”, 1st IAA Planetary Defense Conference, Granada, Spain, 27-30 April 2009. [4] Sanchez, P., Colombo, C., Vasile, M., and Radice, G., “Multicriteria Comparison 360 Among Several Mitigation Strategies for Dangerous Near- Earth Ob-

jects,” Jour- 361 nal of Guidance, Control, and D nam-ics, Vol. 32, No. 1, 2009, pp. 121V142. doi:10.2514/1.36774 [5] Lu, E. and Love, S. “Gravitational Tractor for Tow-ing Asteroids”, Nature, 438, No. 7065, 2005, pp. 177-178. [6] Olympio, J.T., “Optimal Control of Gravity-Tractor Spacecraft 389 for Asteroid Deflec- 390 tion”, Journal of Guidance Control and Dynamics, 33, 3, May-June 2010-, pp. 615-621. [7] Fahnestock, E.G., Scheeres, D.J., “Dynamic Cha acterization and Stabilization of Large 392 Gravity-Tractor Designs”, Journal of Guidance Control and Dynamics, 31, 3, May-June 2008, pp. 501-621. [8] McInnes, C., “Near Earth Object Orbit Modifica-tion Using Gravitational Coupling, ” Journal of Guid-ance, Control, and Dynamics, Vol. 30, No. 3, 2007, pp. 870-873. doi:10.2514/1.25864



EXOMARS 2016 MISSION ANALYSIS: COASTING, ENTRY, DESCENT AND LANDING

D. Bonetti1 ([email protected]), G. De Zaiacomo

1, G. Blanco Arnao

1, J.L. Cano González

1,

C. Parigini1, I. Pontijas Fuentes

1, A.Pagano

1

S. Portigliotti2 ([email protected]), L. Lorenzoni

3 ([email protected])

1 DEIMOS Space S.L.U., Ronda de Poniente 19, Tres Cantos, 28760, Spain

2 Thales Alenia Space Italia, Italy,

3 European Space Agency (ESA), The Netherlands,

The ExoMars programme is pursued as part of a

broad cooperation between ESA and Roscosmos. This

cooperation foresees two missions within the ExoMars

programme for the 2016 and 2018 launch opportunities

to Mars.

The ExoMars 2016 mission, reaching Mars on Oc-

tober, 19th

2016, is led by ESA and has been success-

fully launched from Baikonur by the Russian launcher

Proton-M on March, 14th

2016. The mission is current-

ly on its route to Mars in its assembly configuration

including the Trace Gas Orbiter (TGO) and the Entry,

Descent, and Landing Demonstrator (EDM, named

Schiaparelli), both supplied by ESA. The TGO scien-

tific mission aims at investigating atmospheric trace

gases: it is expected to begin in December 2017 fol-

lowing an aerobraking phase, and to run for five years.

On October 16th

2016, after 7 months of interplanetary

flight and 3 days before landing on the Mars surface

(Meridiani Planum), Schiaparelli will separate from the

TGO and with its mission it will provide Europe with a

demonstration of the technology for entry, descent and

landing (EDL) on the surface of Mars with a controlled

landing orientation and touchdown velocity.

The 2018 mission of the ExoMars programme in-

cludes a carrier Module and a Mars Rover developed

by ESA, and a Descent Module including a Surface

Platform developed by Roscosmos. The project is is

currently in Phase C/D and it is scheduled to be

launched by Proton in 2018.

DEIMOS Space has been involved in the Exomars

Programme (2016 and 2018 missions) since 2004

providing more than 10 years of technical activities in

the areas of End to End (from launch to landing) Mis-

sion Engineering and GNC.

In autumn 2015, the backup launch window of

2016 mission has been activated postponing the launch

to the period 14th-25th March 2016, replacing the

nominal launch window originally set in January 2016.

This paper presents the Mission Engineering activi-

ties performed by DEIMOS Space in support to Thales

Alenia Space Italia, acting as prime contractor for the

ExoMars2016 Mission. Support is dedicated to the

analysis of the Schiaparelli mission, from separation

from the TGO to landing, for the March 2016 launch.

The analyses presented cover the impact of the switch

to the back-up launch window and initial flight predic-

tions for the current launch day, from multiple aspects:

system margins identification through local entry corri-

dors analyses and 3DoF/6DoF End to End Monte Carlo

campaigns, verification of nominal ESA trajectories

and separation maneuver optimization for landing site

targeting, EDM aerodynamic database inspection and

Flying Qualities Analysis, and TGO-Schiaparelli geo-

metric visibility analyses.

All the analyses rely on DEIMOS Space state of the

art tools for Mission Engineering (PETbox, Planetary

Entry Toolbox [1] and LOTNAV, Low-Thrust Inter-

planetary Navigation Tool) whose results and design

methodology for Atmospheric Flight have been recent-

ly Flight Qualified through the successful ESA IXV

mission [2], in which DEIMOS Space was responsible

of the Mission Analysis and re-entry Guidance and

Control.

References:

[1] Bonetti D. et al (2016) “PETbox: Flight Quali-

fied Tools for Atmospheric Flight”, 6th

ICATT.

[2] Bonetti D. et al (2015) “IXV Mission Analysis

and Flight Mechanics: from design to postflight”,

AIDAA 2015.






MARIUS MISSION: Proposal for an ESA M-class seismology mission to Europa. - POSTER M. Van de

Poel1*, M. R. Navarro2, M. Koning3, J Cheung4 and B. Vermeesen5 , 1MSc. student Space Engineering at Delft Uni-

versity of Technology, [email protected], 2,4MSc. student Space Exploration at Delft University

of Technology, 3MSc. student Remote Sensing at Delft University of Technology, 5Prof. Planetary Exploration at

Delft University of Technology

Tidal forces produced by Jupiter cause high stress-

es in Europa's ice shelf, resulting in the cracks and

ridges observed on its surface [1]. Tensile cracks will

open potentially on a diurnal basis caused by the diur-

nal tidal cycle. Due to non-synchronous rotation, ener-

getic Big Bang cracking events, with cracks penetrat-

ing more than 150m deep are expected take place [2].

An ESA M-class mission is proposed by four aero-

space and remote sensing graduate students at the Delft

University of Technology to characterize the Europa

ice shelf and subsurface ocean, based on measuring

seismic waves caused by these Big Bang cracking

events.

The ice shelf thickness will be obtained with a 10%

accuracy by the detection of Crary waves [3], the

thermal structure of the ice shelf can be determined by

measuring the Rayleigh wave dispersion [3][4] and the

ocean depth is measured with an accuracy of 10% by

detecting reflected SCS, PCP and PCS waves [3].

The mission proposal includes a Europa orbiter

spacecraft and two identical battery powered penetra-

tors. The 20.7 kg penetrators carry a three axis seismo-

graph, descent camera and radioscience package to

accurately determine its impact location. The seismo-

meter will measure seismic activity over a one month

period.

[1] Greenberg et al. (1998) Tectonic Processes on Eu-

ropa : Tidal Stresses, Mechanical Response, and Visi-

ble Features. [2] Panning et al. (2007) Long-period

seismology on Europa : 2 . Predicted seismic response.

[3] Lee et al. (2003) Probing Europa' s interior with

natural sound sources. [4] Kovach (2001) Seismic De-

tectability of a Subsurface Ocean on Europa.



Kentucky Reentry Universal Payload System (KRUPS): Design Integration and Overview. C. B. Meek1 and J. M. Cooper1, 1University of Kentucky, Lexington, KY 40506

Introduction: The Kentucky Reentry Universal

Payload System (KRUPS) is a comprehensive tool for collecting atmospheric reentry data of a device using an ablative TPS. The device is an ongoing multidisci-plinary project developed by students with guidance from resident faculty. The ultimate goal of the project is to eject a cluster of KRUPS devices from a high orbital velocity craft in a relatively small space over a short time period. As a result of the project’s highly matrixed requirements and interdisciplinary involve-ment, a systems engineering approach is being imple-mented. Although KRUPS is in its third year of devel-opment, there is still much work to be completed over the next two years. All individual subsystems are either nearly complete or in development. A requirements definition for KRUPS is in progress, along with a ru-dimentary test plan.

Overview: The primary objective for the capsule in 2018 is to complete a mission where the device enters Earth’s atmosphere and transmits data. The capsule will be included in the payload for an ISS resupply mission. It will be ejected from the ISS with the Pacific Ocean as its intended target. The capsule will de-orbit and destructively enter the atmosphere with a ballistic trajectory. During reentry, the device will collect and transmit data through the Iridium satellite constellation before splashdown.

Risk Reduction: In order to mitigate the risks as-sociated with the mission, sequential escalating tests will be performed over the next two years. The first step is a weather balloon launch in April 2016, fol-lowed by a sounding rocket launch in 2017. The bal-loon launch will test data collection, mechanism relia-bility and biological samples. The sounding rocket will test communications, mechanical systems, organic specimen, and data collection in launch and micrograv-ity conditions. These tests, along with implementation of proven technology will help ensure successful col-lection of useful data.

Current Work: The data obtained from this mis-sion will be used to provide insight into environmental conditions during hypersonic reentry of a blunt nosed 45º sphere-cone with an ablative heat shield. This data will be used to help validate various aerothermody-namic codes developed by the University of Ken-tucky’s Gas Surface Interactions Lab.

Because of the interaction of each subsystem with the rest, consistent communication between current development groups is necessary. The previous work which has been done by mechanical engineering senior

design groups is concept development, trajectory anal-ysis, housing construction, Arduino telemetry and data collection, and component selection. Current work is focused on heat shield selection, sizing, instrumenta-tion, construction, and analysis; performed by a me-chanical engineering senior design group. Future work includes code optimization and communications, per-formed by joint electrical engineering and computer science senior design groups.

Future Work: The termination of involvement from the current students will be after the single probe launch from the Space Station. Beyond this, several low cost probes will be launched to establish a strong statistical data set. KRUPS is also being designed for modularity. In the long term, the KRUPS platform can be used to validate new technology in real space condi-tions in a low cost and reliable manner.

Summary: KRUPS is intended to be an inexpen-sive platform for atmospheric data collection during reentry into an atmosphere from space or high altitude. Because of the scope of the project and many interde-pendent subsystems, a systems engineering approach is being employed to effectively and successfully com-plete the single capsule ISS reentry mission. The sys-tems engineering is being performed by a student un-der the advisement of Dr. Suzanne Smith, Systems Engineer/Master Dynamicist and Dr. Alexandre Mar-tin, Fluids Virtuoso. Once this mission is accom-plished, and KRUPS is verified as technically ready, multiple KRUPS devices will be launched at once to obtain a strong statistical mean.

Risk is mitigated by several stages of testing which test new technology, built upon existing technology which has been proven in similar missions. The reentry conditions for design criteria can be simulated by high-ly sophisticated code developed by the Gas Surface Interactions Lab. The shape of the capsule, based on REBR1, was flight qualified in 1976 at NASA Langley Research Center2. This, along with other technology, provides a sound base for a reliable reentry test plat-form.

References: 1Weaver, M. A. and Ailor, W. H., “Reentry

Breakup Recorder: Concept, Testing, Moving For-ward,” AIAA Space 2012

2Brooks, J. D., “Some Anomalies Observed in Wind Tunnel Tests of a Blunt Body at Transonic and Supersonic Speeds,” NASA TN D-8237, June 1976.



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RAPID EXPLORATION OF THE SOLAR SYSTEM USING AEROGRAVITY-ASSIST. P. J. Edelman1†, E., Shibata2†, and S. J. Saikia3†, J. M. Longuski4†, 1Ph. D. Candidate, [email protected], 2Graduate Student, [email protected], 3Visiting Assistant Professor, [email protected], 4Professor, [email protected], †School of Aeronautics and Astronautics, Purdue University, 701 W. Stadium Ave., West Lafayette, IN, 47907-2045.

Introduction: Aerogravity-Assist (AGA) maneu-

vers are a high risk, but high reward augmentation of traditional planetary flybys. A spacecraft intentionally travels through a planet’s (or moon’s) atmosphere dur-

ing the flyby and harnesses the additional aerodynamic forces to increase the total amount of turning that can be achieved. In a normal gravity-assist, the amount of turn-ing is limited by the height of the atmosphere, as enter-ing reduces the orbital energy or the mission would likely fail without an aerodynamically inclined vehicle (see Fig. 1). The flyby limitation in turn restricts the number of potential targets in the solar system, forcing mission designers to wait until planetary alignments are suitable for a chosen mission. Using an AGA maneuver, the turn angle can be arbitrarily chosen, provided enough orbital energy remains and the spacecraft can withstand the physical conditions such as aerodynamic heating, g-loading, and dynamic pressure. With the turn angle now able to be increased, the design space for in-terplanetary missions is expanded, increasing the num-ber of available launch dates, and reducing the total mis-sion time-of-flight (TOF).

Fig. 1: Comparison between AGA and a traditional gravity-assist, adapted from [1,12].

The potential applications for AGA are numerous and can deliver a wide variety of scientific instruments. Previously proposed missions using AGA have in-cluded fast atmospheric science and sample return mis-sions to both Venus and Mars with about one year flight times [2] and missions to the outer planets with a se-verely reduced TOF [3]. All of the inner planets (with atmospheres) are good candidates for AGA. Mars is es-pecially intriguing for AGA, as its gravity-well cannot often provide a desirable heliocentric energy change when the spacecraft approaches at high relative speeds. The AGA maneuver essentially boosts the gravity-assist to mimic the result as if it were performing the gravity-

assist around a larger planet such as Earth or Venus. An-other promising candidate is Saturn’s moon Titan. An

AGA at Titan has the potential to assist in capture around Saturn (an aerocapture maneuver at Saturn via Titan), or to assist in getting to other targets in Saturn’s

orbit. A program has been developed at Jet Propulsion La-

boratory, and later became automated at Purdue Univer-sity, called STOUR-AGA (Satellite TOUR design pro-gram with AGA) [2]. The program finds all interplane-tary trajectories within a set of user-specified mission parameters such as flyby sequence, launch dates, time-of-flight, launch speeds, and entry speeds. The AGA portion is approximated using an analytical approach where the vehicle is assumed to fly at a constant altitude and lift-to-drag ratio (L/D). Next the program optimizes the AGA maneuver by minimizing the total heat load the vehicle accrues for a given set of aerodynamic char-acteristics. Total heat load is the metric being minimized because a large percent of the mass will go towards ther-mal protection from the atmosphere at hypersonic speeds. If the trajectory flown mitigates the heat load, more mass can be allocated towards payload and scien-tific instruments.

Figure 2 represents a 458 day Earth-Venus-Mars-Earth trajectory (using STOUR-AGA program) with AGA at Mars only. The Earth launch, Venus and Mars encounter, and Earth return dates are labeled in the fig-ure. This type of trajectory would not be feasible using traditional gravity assists. The total hyperbolic turn an-gle at Mars using AGA is 72.6°, while a traditional grav-ity assist with a distance of closest approach of 200 km (in order to avoid the atmosphere) only yields a turn an-gle of 27.3°. Additionally, there is about a 120-day re-duction in TOF when comparing the presented trajec-tory with one not using AGA, as found by Hughes et al. [4]. However, AGA trajectory also has the advantage of passing through the atmosphere, thereby acting as an at-mosphere probe, not possible for a traditional gravity assist spacecraft.

The type of vehicle most suitable for AGA is the aer-odynamic waverider, producing the highest L/D at hy-personic speeds [5]. The Boeing X-51 was a mission re-cently flown, demonstrating the capabilities of the wa-verider design. Waveriders take advantage of the at-tached shock the body produces by having the leading edge ‘ride’ the shock wave to increase the L/D. From a flowfield perspective, the leading edge maintains a pres-



sure difference between the top and bottom surfaces, al-lowing the higher pressure from the lower surface con-tribute to the vehicle’s lift. The increased lift is present

even at zero angle-of-attack, whereas a sharp hyper-sonic vehicle must maintain an angle of attack to pro-duce lift [6].

Fig. 2: A sample fast (458) Earth-Venus-Mars-Earth

(EVMA) trajectory with aerogravity-assist at Mars and corre-sponding encounter dates.

One defining characteristic of waverider design is that it is an inverse design process [6]. As the waverider is dependent upon the shock it creates, the flowfield is generated first, then the vehicle is shaped to create the flow. The flowfield itself is derived from an Euler flow about an infinitely long wedge- or conical-based shape. Although the basic flowfield would be a right frustum cone [6], other shapes can be used for this purpose, in-cluding osculating cones [7] and crosses [8]. Surfaces of the waverider come from streamsurfaces of the inviscid flowfield being truncated at an appropriate length, with the leading edge connecting the upper and lower surface and touching the shock wave. Other factors, such as thermal energy constraints and viscous flow, also adjust the shape by either blunting the leading edge [9] or smoothing the upper and lower surfaces [10].

Proposed AGA Mission Concept: Using AGA pre-sents a paradigm-shift in planetary exploration in terms of vehicle design. Since a waverider has never flown on an interplanetary mission, careful design considerations need to be made to accommodate desired scientific ob-jectives. Traditionally, scientific spacecraft tend to be bulky, with equipment such as solar panels and antennas protruding at different angles from the main body. For a waverider, all equipment must be designed to fit within the aerodynamic shell of the vehicle or able to retract, so they do not become damaged during the AGA ma-neuver. Figure 3 shows a traditional interplanetary spacecraft (Cassini orbiter with the Huygens entry probe) vis-à-vis a Waverider (based on the X-51 de-

sign). We observe that AGA mission will require a re-think of spacecraft design using traditional knowledge. In this paper, we will present a mission architecture for the rapid exploration of two planets (Mars and Venus) using aerogravity assist trajectories. The work includes end-to-end conceptual mission design which includes design of AGA spacecraft, selection and accommoda-tion of instruments, and design of spacecraft subsys-tems.

Fig. 3: On the left is a traditional spacecraft (Cassini-Huygens, adapted from JPL [11]), while on the right is a waverider (Boe-ing X-51, adapted from [12]).

Benefits of AGA Missions: There are several bene-fits of AGA missions: AGA, clearly reaps the benefits of both traditional

gravity-assist as well as atmospheric passage. Since AGA mission is significantly shorter than tra-

ditional missions, cost can be saved due to reduced mission operations cost.

An AGA spacecraft can be used to return atmospheric sample from planets.

An AGA mission, such as the EVMA, allows the rapid exploration of two planets. At the two encoun-ters, miniature atmospheric probes can be released to obtain directed yet high-quality science data.

Due to relatively shorter mission time, scientists can get science data quickly compared to long waits for conventional missions.

References: [1] Longuski J. M. (1982) Jet Propul-sion Laboratory, EM 312/82-133. [2] Henning G. A. et al. (2014) J. of Spacecraft and Rockets (JSR), Vol. 51, No. 6, 1849-1856. [3] Sims J. A. et al. (2000) JSR, Vol. 37, No. 1. [4] Hughes et al. (2015) JSR, Vol 52, No. 6, 1712-1736. [5] Anderson J. D. et al. (1991) JSR, Vol. 28, No. 4. [6] Anderson J. D. et al. (1991) 29th Aero-space Sciences Meeting, AIAA 91-0530. [7] Takashima N. and Lewis M. J. (1999) J. of Aircraft, Vol. 36, No. 1. [8] Kai C. et al. (2007) Acta Mechanica Sinica, Vol. 23, 247-255. [9] Armellin R. et al. (2007) JSR, Vol. 44, No. 5. [10] Bowcutt K. G. et al. (1987) AIAA 25th Aerospace Sciences Meeting AIAA-87-0272. [11] JPL, http://www.jpl.nasa.gov [12] USAF, http://www.af.mil



ROBOTIC GRIPPER DESIGN AND TESTING FOR POTENTIAL MARS SAMPLE RETURN (MSR).

R. L. McCormick1, E. A. Kulczycki2, D. Wai3, L. N. DuCharme4 and T. Srinivasan5 1NASA JPL (mccor-

[email protected]), 2NASA JPL ([email protected]), 3NASA JPL ([email protected]), 4NASA JPL

([email protected]), 5NASA JPL ([email protected]).

Introduction: Mars Sample Return (MSR) was

identified by the 2011 planetary science decadal survey

as a high priority long-term goal [1] for NASA. A

three-mission campaign concept is currently being in-

vestigated. A second Sample Return Lander (SRL)

mission concept could potentially collect the sample

tubes left by M2020—or another rover—and load them

into a Mars Ascent Vehicle (MAV) to be launched into

Mars orbit.

Previous efforts have demonstrated the feasibility

of picking a sample tube from a flat, sandy surface,

reorienting the tubes for axial manipulation [2], and

inserting the tubes into a container [3] for launch into

Mars orbit. Mobility technology required for driving

to sample tube locations has been demonstrated on

Mars Science Laboratory (MSL) [4-6]. A JPL internal

research task aims to study and test methods of reliably

acquiring sample tubes from the ground and placing

them into the MAV. This paper focuses on the gripper

design for the ground pickup portion of the sample

transfer chain.

Gripper Design: A baseline strategy for this task

involves grabbing the body radially from the ground.

This grip eliminates the issue of the end-effector and

arm volume coming into contact with the ground.

Based on the uncertainty of positioning the MSL robot-

ic arm, an uncertainty of end effector placement of +/-

10mm in the vertical, lateral, and axial direction was

used. The estimated acceptable position uncertainty is

demonstrated in Fig. 1. A potential flight mission

could redesign the arm to be optimized for to pick up

and potentially reduce the positioning uncertainty.

An active pair, two finger impactive gripping strat-

egy was chosen for tube prehension testing. Based on

initial research and testing, a two-finger, two point con-

tact gripping method [7] provided reliable, secure grip.

This grip method, as shown in Fig. 2, is similar to a test

tube holder.

This strategy provides a simple, robust method for

tube acquisition. A custom parallel jaw gripper, shown

in Fig. 3, was designed and built to allow for increased

flexibility and modularity. Custom gripper jaws, link-

ages, and accessories could quickly developed and

integrated with the custom gripper.

A custom robotic arm was designed and built to

further study picking up samples on Mars. This arm is

shown in Fig. 4. The robotic arm is similar to the 2 m

length of Curiosity rover arm, but has an additional

degree of freedom (DOF). An additional DOF was

added to enable gripper roll.

Each gripper design was tested in the Planetary Ro-

botics Laboratory located at NASA Jet Propulsion La-

boratory located in Pasadena, CA. Results of this test-

ing led to gripper design iterations. Over 150 test runs

were performed during this phase of development.

Testing scenarios included level flat rock, local mini-

ma on a concave rock surface, angled contact with rock

and sand, local minima between two rocks, tube under

Fig. 1: Laterial and axial (Left) and vertical (Right)

uncertainty for succesful body grip ground pickup

Fig. 2: Two finger, two contact gripper method

used for tube prehension

Fig. 3: Custom 2-finger parallel gripper mechanism

Fig. 4: Custom 6 DOF robotic arm in testing area



lip of a rock in contact with sand, tube under lip of

rock in contact with rock, local minima between rocks

(angled contact with rock), tube in “tube sized” crevice

with one side open, tube in “tube size” crevice with

rocks on all sides, two high points on concave rock

surface, wedged in crevice between two rocks. Rego-

lith testing scenarios include on the surface, buried at

12 mm, 24 mm, and 38 mm at varied angles. The test-

ing regolith types include dust, sand, and a mixture of

sand and dust.

Harpy Eagle. While the dual purpose grippers used

in the demo performed well for the rapid demonstration

task [2], the size and shape of the sample tube had sig-

nificantly evolved. To accommodate the tube change,

the gripper size was scaled up and machined out of

aluminum instead of rapid prototyped PLA plastic ma-

terial. The resulting gripper design is shown in Fig. 1.

Of the 43 ground pickup tests, 35 were successful.

Marsh Hawk. The Marsh Hawk grippers use the

basic gripping and positioning uncertainty accommoda-

tion strategy as the Harpy Eagle while aiming to im-

prove upon the shortcomings of the Harpy Eagle. The

new side profile allows the grippers to reach between

tighter areas for rock cases, such as in a local minima

of a concave rock surface, or between rocks, as shown

in Fig. 5. The lower front profile reduces regolith

buildup for pickup on or buried in sand. Although the

lower profile bottom surface was intended to reduce

force required to reach buried tubes at various depths,

testing did not reveal a significant reduction. A graph

comparing force required to reach various depths in

regolith is shown in Fig. 6. During testing of the

Marsh Hawk grippers, a force control, terrain follow-

ing technique was developed to assist in reducing ver-

tical uncertainty on hard surfaces. Of the 75 ground

pickup tests, 63 were successful.

Compact Bar. The Marsh Hawk grippers were rede-

signed as a way to both reduce vertical uncertainty and

to reduce the width of the grippers by adding cross

bars. The grippers could be move down vertically until

contact is read by the force torque sensor, as shown in

Fig. 7. These grippers are designed to be agnostic to

the surface type. If the bars come down into contact

with regolith, they'll continue on until the bars to come

contact the tube. If the jaws come in first contact with

a hard surface, the vertical position of the tube resting

on the hard surface is also known. By reducing the

vertical uncertainty, it is possible to reduce the length

and width of the gripper profile. This allows the grip-

pers to reach into tighter areas, such as small rock crev-

ices. While testing is in the initial phases, of the 24

ground pickup tests, 24 were successful.

Conclusion: The work presented in this paper

demonstrates the feasibility of acquiring sample tubes

from various Martian terrain scenarios. This testing

led to a list of recommendations for the Mars 2020

rover mission, including required tube feature design

changes and tube drop off terrains. Future work of this

research task will focus on improved ground acquisi-

tion and autonomy.

Acknowledgment: The research described in this

publication was carried out at the Jet Propulsion La-

boratory, California Institute of Technology, under a

contract with the National Aeronautics and Space Ad-

ministration (NASA). Copyright 2016 California Insti-

tute of Technology. U.S. Government sponsorship

acknowledged.

[1] R. Mattingly and L. May (2011) AeroConf, 1-13.

[2] K. Edelberg et al. (2015) AIAA SPACE, 1-10. [3]

P. Backes et al. (2014) AeroConf, 1-9. [4] M. W.

Maimone et al. (2007) ICRA. [5] R. A. Lindemann et

al. (2006) IEEE Robot. Autom. Mag., vol. 13, no. 2,

19-26. [6] M. Bajracharya et al. (2008) Computer, vol.

41, no. 12, 44-50. [7] G. J. Monkman et al. (2007) Ro-

bot Grippers.

Fig. 5: Size comparison between Harpy Eagle (red)

and Marsh Hawk (blue) grippers

Fig. 6: Force plots for vertical sand insertion at var-

ious depths for Harpy Eagle and Marsh Hawk

grippers

Fig. 7: Compact Bar grippers contacting tube to

provide reduce vertical uncertainty before closing



SATURN-URANUS TRAJECTORIES FOR MULTI-PLANET MISSIONS. K. M. Hughes1, P. Agrawal

2, G. A.

Allen2, A. J. Mudek

3, S. J. Saikia

3, J. M. Longuski

3, H. H. Hwang

2, and E.Venkatapathy

2,

1Purdue University School

of Aeronautics and Astronautics, Armstrong Hall, 701 W. Stadium Ave., West Lafayette, IN, 47907,

[email protected], 2NASA Ames Research Center, Moffett Field, CA 94035,

3Purdue University School of

Aeronautics and Astronautics, Armstrong Hall, West Lafayette, IN, 47907.

Introduction: In 1986, the Voyager 2 spacecraft

achieved the only close encounter with Uranus, fol-

lowed in 1989 by a flyby of Neptune. These two Ice

Giants remain among the least known planets of our

Solar System, and are the only class of planet never to

have been explored in detail. At the August 2015 meet-

ing of the Outer Planets Assesment Group, NASA’s

Planetary Science Division announced the initiation of

an Ice Giants Study for a Flagship-class mission. The

JPL-lead study is to be part of the next Decadal Sur-

vey—beyond 2022 [1].

The announcement of the Ice Giants Study speci-

fied, as part of it’s objectives, to identify enabling

technologies, and assess the capabilities afforded by

the SLS [1]. This work identifies mission opportunities

to Saturn and Uranus, which would be enabled with the

use of the SLS Block 1B. This mission would not only

send a probe and orbiter to Uranus, but also a probe to

Saturn.

Methods: In this study, we investigate multiple

gravity-assist, ballistic trajectories to Uranus that ex-

plore 26 combinations of gravity-assist paths. All of

these paths include Saturn, but some also include gravi-

ty assists at Venus, Earth, Mars, and Jupiter. We search

for trajectories with launch dates spanning from Janu-

ary 2023 to December 2038, using a patched-conic

model with an analytic planetary ephemeris, via the

program STOUR [2]. Solutions capable of delivering a

payload of at least 2000 kg become available when the

Space Launch System (SLS) launch vehicle is em-

ployed, and are estimated to be sufficient for a multi-

planet, Saturn-Uranus mission. Such a mission, for

example, could send a probe to Saturn on the way to

delivering a probe and orbiter to Uranus.

A set of possible approach trajectories are modeled

at the Saturn and Uranus encounters, to assess feasibil-

ity for probe entry, and what (if any) adjustments are

needed for ring avoidance. Gravity-assist trajectories of

interest are optimized using the Mission Analysis Low-

Thrust Optimization (MALTO) tool [3].

Key Findings and Conclusions: The trajectory

results from the STOUR search are sorted by total

flight time and delivered mass at Uranus. The deliverd

mass estimate for each trajectory is computed assuming

an Isp of 330 sec, and capture into a 20-day elliptical

orbit about Uranus, with periapsis radius of 52,000 km.

A %13 mass penalty of the total propellant mass is also

included in the delivered mass calculation, to estimate

the mas of the propellant tanks.

Among the thousands of results from this prelimi-

nary study, we found several trajectories of interest

between 2023 and 2038. One such trajectory, with

lauch date on September 13, 2028, can deliver 2,200

kg to Uranus in 10.5 years, using a Delta-V Earth grav-

ity assist, followed by a Saturn gravity assist. This tra-

jectory, shown in Fig. 1, was found using MALTO,

based on an initial guess found by STOUR. The deliv-

ered payload at Uranus should provide sufficient mass

for a unique two-planet mission; delivering, a probe to

the Saturnian system in 5.3 years, followed by a 5.2-

year trip to Uranus, with a second probe and an orbiter.

Figure 1. 2028 Saturn-Uranus opportunity.

The SLS will change how we think about what is

feasible for future space exploration missions, and

these Saturn-Uranus opportunities highlight one such

game-changing mission concept.

References: [1]Green, J. L. (2015), “Planetary Sci-

ence Division Satus Report,” OPAG, Aug. 24 Presenta-

tion. [2]Longuski, J. M. and Williams, S. N., (1991)

Celest. Mech. Dyn. Astr., 52, 207–220. [3]Sims, J. A.

et al. (2006) AIAA/AAS ASC, #2006-6746.



PLANETARY PROBE ENTRY ATMOSPHERE RECONSTRUCTION USING SYNTHETIC AIR DATA SYSTEM. C. D. Karlgaard1 and M. Schoenenberger2, 1Analytical Mechanics Associates, Inc., 2Atmospheric Flight and Entry Systems Branch, NASA Langley Research Center.

Introduction: An atmospheric state estimator is developed, based on inertial acceleration and angular rate measurements combined with an assumed vehicle aerodynamic model. The approach utilizes the full nav-igation state of the vehicle to recast the vehicle aerody-namic model to be a function solely of the atmospheric state. Force and moment measurements are based on vehicle sensed accelerations and angular rates. These measurements are combined with an aerodynamic model and a nonlinear least squares approach is uti-lized to estimate the atmospheric conditions.

The approach differs from previous aerodynamic model-derived atmospheric estimation methods in that the algorithm is based on a stochastic rather than a completely deterministic method. The potential bene-fits of the new method are that the input measurements are weighted appropriately, optimal estimates are com-puted based on these weights, and output statistics are readily available. Moreover, the approach includes estimates of atmospheric winds directly in the filter without a secondary post-processing method, and the algorithm is simple enough that onboard implementa-tion is a possibility.

Application to Mars Science Laboratory: On August 5th 2012, the Mars Science Laboratory (MSL) entry vehicle successfully entered the atmosphere of Mars and landed the Curiosity rover safely on the sur-face of the planet in Gale crater. The MSL entry vehi-cle was comprised of a 70-degree sphere-cone heat-shield and backshell consisting of a stack of three trun-cated cones. The forebody was similar to the geometry developed for the Viking Mars landers.

MSL carried with it an instrumentation package designed to measure the aerodynamic and aerothermal environments during atmospheric entry. This instru-mentation package was known as the MSL Entry, De-scent, and Landing Instrumentation (MEDLI) [1], which consisted of three major subsystems: the Mars Entry Atmospheric Data System (MEADS), the MEDLI Integrated Sensor Plugs (MISP), and the Sen-sor Support Electronics (SSE). The MEADS consisted of seven pressure transducers connected to flush orific-es in the heat shield to measure pressures across the vehicle forebody. The MISP devices were a system of seven thermocouple and recession sensors that provid-ed aerothermal measurements of the heat shield per-formance. The SSE provided power to the sensors, conditioned their signals, and transmitted the data to storage on the Curiosity rover. The MEDLI flight data can be used as a validation case to compare the results

of a synthetic air data system with those computed from the MEADS pressures.

An example result is shown in Figure 1. This figure compares the reconstructed capsule angle of attack from the MEADS pressure data [2] and the results of the new syntheric air data system algorithm. The two algorithms produce similar reconstructed time histo-ries. Estimates of the northly component of atmospher-ic winds are shown in Figure 2. The estimates are gain similar, and are consistent with observed winds de-duced by the vehicle guidance response [3]. Similar results are also found in other reconstructed air data parameters such as density, dynamic pressure, and Mach number. These comparison indicates that the new method provides estimates consistent with the air data measurements, without the use of pressure data.

Figure 1 – Angle of Attack Reconstruction

Figure 2 - North Wind Reconstruction

References: [1] Cheatwood, F. H., et al (2014) NASA TM-

2014-218533. [2] Karlgaard, C. D. et al. (2014) JSR, 51, pp. 1029-1047. [3] Chen, A. et al, JSR, 51, pp. 1062-1075.



GEER: A unique capability to support planetary exploration and probe development T. Kremic1, D. Vento1,

L. Nagley1, NASA Glenn Research Center 21000 Brookpark Rd, Cleveland, Ohio 44135. ([email protected]).

Introduction: As scientists continue to push the

frontiers of exploration and knowledge of our solar

system, more complicated and challenging missions

will be conceived and need to be successfully imple-

mented. Those in the planetary science community

know that the trend in planetary exploration has been to

observe remotely and then move closer on subsequent

missions to the particular destination with the goal of

eventually acquiring and returning samples to Earth.

To successfully implement this exploration approach

demands an increasing ability to understand and simu-

late the environments that one seeks to explore [1, 2].

The Glenn Extreme Enviornment Rig (GEER) is a re-

cent addition to the planaetary science arsenal that

helps the science and mission communities understand

and prepare for these challenging in-situ missions.

GEER:

GEER provides a unique capability to simulate at-

mospheric chemistry of virtually any body in the solar

system even the Venusian atmosphere, including pres-

sure and temperature conditions down to its surface.

GEER was conceived to be an asset for science, tech-

nology, and planetary mission communities that have a

growing interest in the unique physical characteristics

and processes that occur extreme conditions such as the

surface of Venus or deep in the Satern atmposhere. In

addition to meeting needs of science and experiment-

ers, GEER supports the needs of technologists devel-

oping instruments and hardware at the component or

system level. Probe mission planners can also benefit

from GEER through risk reduction tests, experiments

to verify measurement techniques or help interpret fu-

ture results.

This poster presents the features and characteristics of

the NASA’s Glenn Extreme Environment Rig (GEER),

which is located in Cleveland, Ohio. In addition to

providing basic features the poster discusses some of

the knowledge and questions that are being or can be

explored with GEER’s capabilities. The basic opera-

tions approaches and constraints are described as is

basic information for prospective users.

Figure 1. Schematic view of the GEER operation sys-

tem.

References: [1] Venus Exploration Analysis Group

(VEXAG), “Venus exploration themes” February

(2014). [2] Venus Exploration Analysis Group

(VEXAG), “Roadmap for Venus exploration” March

(2014).




Hypervelocity Expansion Tube Studies of Blunt Body Aerothermodynamics in CO2. M. G. Leibowitz1, B. R. Hollis2, and J. M. Austin1, 1California Institute of Technology, Pasadena, Ca 91125, 2NASA Langley Research Center, Hampton, VA 23681.

Introduction:

To be capable of launching higher mass vehicles for robotic sample return and human exploration to Mars, improved accuracy in the prediction of aerodynamics and thermal loads must be made. Problematically, cur-rent state-of-the-art simulations are limited by a lack of experimental data where the free-stream thermochemi-cal state is not well characterized. If the flow is accel-erated from stagnation to hypersonic conditions through a nozzle, freezing can occur and the freestream will be partially dissociated. (Nompelis et al. 2003, 2007). The entire nozzle flow would have to be simu-lated and the freestream predictions of today are deemed to be inadequate. (Nompelis et al. 2010, Doraiwamy et al. 2010).

To overcome this challenge, the proposed experiments will be carried out in an expansion tube facility where freestream dissociation is minimized, allowing for an improved database of surface and flow field measure-ments at hypervelocity conditions. The Hypervelocity Expansion Tube (HET) facility was specifically built to eliminate the experimental challenge of freestream dissociation by accelerating the test gas through an unsteady expansion wave, rather than a nozzle. The HET has the capability of reaching a broad range of test gas conditions simply by changing the initial pres-sure and gas composition in the facility sections.

A laminar heat transfer database (Hollis et al. 2005) for blunt bodies in a non-dissociated freestream in CO2 and air will be obtained using thermocouple. The thermocouples gauges that are used in these experi-ments will be type E coaxial thermocouples, made using the design of Sanderson (1995). They have been proven to have an adequate signal and are robust and survive the hypersonic conditions in the HET. These heat transfer measurements will provide additional data for physics-based model development.

Additionally, time-resolved high speed schlieren imag-ing will be used to examine the geometries of NASA interest. This is needed because the current State-of- the-art simulations show poor agreement in high en-thalpy (≥5 MJ/kg) CO2 flows with high thermochemi-cal activity.

The widely known shock shape shows a response to non-ideal freestream thermochemistry. The experi-mental capability known at our facility can match the enthalpies and velocities where anomalous results were reported and produce MSL shock shapes that are in good agreement with simulations tested at HET condi-tions.

This work will improve the entry modeling and simu-lation models for a hypersonic, chemically reacting environment. The intent to launch larger-scale vehicles requires the anomalies in the uncertainty presented to be resolved.

References: [1] Doraiswamy, S., Kelley, J., and Candler, G.V.,

Vibrational Modeling of CO2 in High- Enthalpy Noz-zle Flows, J.Therm. Heat Transfer 24 (1):9-17, 2010.

[2] Hollis, B. R., Leichy, D. S., Wright, M. J., Holden, M. S., Wadhams, T. P., MacLean, M., and Dyakonov, A., AIAA Paper 2005-1437, 2005.

[3] Nompelis I., Candler, G.V., and Holden, M.S., Effect of Vibrational Nonequilibrium on Hypersonic Double-Cone Experiment AIAA J. 41 (11):2162-2169, 2003.

[4] Nompelis, I., and Candler, G.V., Investigation of Hypersonic Double-Cone Flow Experiments at High Enthalpy in the LENS Facility, AIAA Paper 2007-203, 2007.

[5] Nompelis, I., Candler, G.V., Numerical Investi-gation of Double-Cone Flow Experiments with High-Enthalpy Effects , AIAA Paper 2010-1283, 2010.

[6] Sanderson, S. R., “Shock Wave Interaction in Hypervelocity Flow,” Ph.D. Thesis, California Inst. of Technology, Pasadena, CA, 1995.



LOW DENSITY SUPERSONIC DECELERATOR (LDSD) SUPERSONIC FLIGHT DYNAMICS TEST

(SFDT) PLUME INDUCED ENVIRONMENT MODELLING B. L. Mobley1, S. D. Smith2, J. W. Van Norman3,

S. Muppidi4 and I. Clark5, 1NASA Marshall Space Flight Center, Huntsville, Alabama, 35811, 2Plumetech, Huntsville,

Alabama, 35811,3Analytical Mechanics Associates, Hampton, Virginia, 23666, 4NASA Ames Research Center,

Mountain View, California, 94035, 5NASA Jet Propulsion Laboratory, Pasadena, California, 91109.

Abstract: The methodology and development of

plume induced aerodynamic and heating models for the

Low Density Supersonic Decelerator (LDSD) test vehi-

cle is presented. The LDSD test vehicle falls within a

unique class of balloon-assisted launch vehicles whose

specific purpose is to test supersonic deceleration tech-

nologies (suited for Mars payloads) within the Earth’s

upper stratosphere. To achieve freesteam test conditions

similar to that experienced during Mars entry, the vehi-

cle is tethered to a helium balloon to ascend to an alti-

tude of approximately 36.5 km (120 kft) where it is re-

leased and is accelerated along a ballistic trajectory. A

single solid rocket kick motor burns for approximately

65 seconds throughout the powered phase, lofting the

vehicle to approximately Mach 4.2 at 53 km (170 Kft).

Vehicle attitude spin stabilization is initiated prior to as-

cent using two pairs of solid rocket motors. Post pow-

ered phase, an additional two pairs of solid rocket mo-

tors are fired to terminate the spin. Two Supersonic

Flight Dynamics Tests (SFDT-1 and SFDT-2) occurred

on June 28, 2014 and June 8, 2015, respectively. Pre-

flight and post-flight plume induced aerodynamic and

heating models were generated to predict the plume in-

duced forces, moments and heat fluxes experienced dur-

ing the spin-up, powered ascent, and spin-down phases.

Comparisons of the predictions with post-flight recon-

struction data are provided. Relatively good agreement

was observed between the flight data and predictions

from the highest fidelity models.



SIZING METHODS FOR ADVANCED MARS ENTRY DESCENT AND LANDING SYSTEMS. M. A. Lob-

bia1, 1Jet Propulsion Laboratory, California Institute of Technology (4800 Oak Grove Drive, MS 301-490, Pasadena,

CA 91109; [email protected]).

Introduction: Mars provides a unique environ-

ment for entry, descent and landing (EDL) missions. It

is frequently argued that the atmosphere is thick

enough that an engineer must worry about the aerody-

namic and aerothermal aspects of entry, yet not dense

enough to provide the required deceleration to land

large payloads to the surface. This can lead to novel

EDL approaches such as the Mars Science Laboratory

(MSL) Skycrane system. [1] For future Mars missions

with higher-mass payloads, however, advanced EDL

techniques such as Supersonic Retro-Propulsion (SRP)

can be considered as one means to reduce the risk in

scaling up supersonic parachutes or relying on addi-

tional deceleration devices such as Hypersonic Inflata-

ble Aerodynamic Decelerators (HIADs). [2]

An essential part of the initial trade space explora-

tion for future Mars mission designs is the mass-sizing

of the entry vehicle. While a detailed Master Equip-

ment List (MEL) provides a reasonable bottoms-up

approach to developing an accurate estimate of the

vehicle components and masses, conceptual design

studies frequently do not provide enough definition of

the vehicle configuration to make this possible. In

many cases, a first-order mass estimate is generated

using simple scaling relationships, such as sizing the

entire heatshield mass based on the entry heat load for

the vehicle. The drawback to the techniques, however,

is that in many cases these empirical relationships are

not applicable to configurations that deviate substan-

tially from heritage designs.

Objectives and Modeling Approach: The present

work discusses the development of a modular frame-

work for mass-sizing of planetary entry systems, with

an initial focus on Mars missions. The framework is

designed to bridge the low-fidelity simple scaling rela-

tionships discussed above by allowing the use of phys-

ics-based sizing methods where possible; it also in-

cludes an automatically-generated MEL that can be

used as a starting point in more detailed follow-on

studies.

The framework divides the vehicle into various el-

ements, which currently includes: Aeroshell (e.g. heat-

shield and backshell), Aerodynamic Decelerator (e.g.,

supersonic parachute), Ejectable Masses, Descent

Stage, and Cruise Stage. For each of these elements,

various sub-models are implemented. For example, the

Descent Stage component currently has sub-models for

both a Skycrane system (similar to MSL), as well as a

landed descent stage for SRP applications.

This framework uses a spreadsheet format to allow

rapid incorporation of new sub-models, as well as easy

updates and refinements of existing sizing methods. A

programming-like approach is utilized [3], where a

“Main” worksheet has the primary model inputs, and

each major element of the system (e.g. Aeroshell, De-

scent Stage) is on a separate worksheet. For each of

these element worksheets, various sub-models are

listed, and automated logic is used to link the correct

sub-model into the integrated sizing based on the input

selections on the Main worksheet.

An iterative approach is required to size the entry

mission – for example, the capsule backshell structure

mass is calculated as a function of the vehicle entry

mass, which is a function of the backshell mass. There-

fore, an optimization approach is implemented to iter-

ate specific sizing variables as needed. This also allows

the use of specific objective functions to tailor the de-

sign to desired requirements (e.g. sizing for a fixed

payload mass, or optimizing the payload mass for a

desired entry ballistic coefficient).

Additional details of the sizing methods for specific

elements and components, as well as some historical

sizing relationships developed, will be discussed in the

final manuscript.

Preliminary Results: The framework was validat-

ed against the MSL mission. Based on the results, it

was observed that the model was able to predict the

entry mass of MSL to within 3%.

Additional results will be provided in the final

manuscript to highlight the application of the model to

potential future Mars mission concepts, such as an SRP

concept to increase landed payload mass to the Martian

surface. Examples of sizing for ballistic coefficient will

also be presented, as will further discussion of TPS

sizing (based on the entry trajectory heating environ-

ments).

References: [1] Steltzner, A. D., et al (2014) Jour-

nal of Spacecraft and Rockets, 51, 4, 994-1003.

[2] Braun, R. D., and Manning, R. M. (2006) IEEEAC

#2006-0076. [3] Lobbia, M. A. (2006) AIAA Paper

2006-821.



SURRENDER, AN IMAGE RENDERING SOFTWARE FOR SCIENTIFIC SPACE SCENE SIMULATION - POSTER. R. Brochard1 and N. Despré1, 1Airbus Defence and Space, 31 rue des Cosmonautes, 31402 Toulouse Cedex, France, ([email protected] , [email protected]).

Introduction: The use of cameras and lidars as

navigation sensors on space vehicles, and in particular on exploration probes, is an emerging topic of the last 15 years. The development and maturation of this field – known as computer vision – for space exploration require a large amount of images, under a variety of conditions. Real images from actual space exploration missions can rarely be used directly for this task [1], as the accuracy of the associated ground truth is most of times insufficient with respect to the needs of comput-er vision algorithms validation. Image simulation is thus a mandatory step for space exploration related computer vision projects. Although several image sim-ulation tools are publicly available, none of them ad-dress fully the specific challenges of scientific image rendering of space scenes. For this reason, Airbus De-fence and Space has been developing since 2011 its own image simulation tool, SurRender. As a scientific rendering software dedicated to space exploration mis-sions, SurRender aims at a high representativeness and physical accuracy.

Figure 1: Examples of space scenes simulated with SurRender, including a depth-map (bottom-left)

Scientific-level image rendering: SurRender’s images are rendered by raytracing. The physical prin-ciples of light propagation (geometrical optics) are implemented within the raytracer. Physical optics-level effects, such as diffraction of the light by the aperture of the imaging instrument, are taken into account at macroscopic level trough the user-specified Point Spread Function (PSF). Interaction of the light with the surface of the scene objects is modelled in terms of Bidirectional Reflectance Distribution Function

(BRDF). The image is rendered in physical units: each pixel contains an irradiance in W/m².

The images generated by SurRender are physically accurate, and are thus suitable for scientific applica-tions, from radiometric studies and sensor design with respect to the expected scene content, to the validation of computer vision algorithms performance.

Figure 2: Jovian moons simulated by SurRender (left) and real image from New Horizon mission. Planetshine on Io is correctly simulated. Note the volcanos around Io, not yet implemented in Sur-Render.

Figure 3: Limb of Ganymede simulated by Sur-Render (left) and as imaged by LORRI (right).

Optimized for space scenes: SurRender is opti-

mized for space scenarios with huge distances between objects. All computation are performed in double pre-cision, which is sufficient to cover numerical precision needs for scenes sizes up to solar system size. Space scenes are sparse: SurRender explicitely takes ad-vantage of this property and is capable of focusing the raytracing directions on energy sources. This drastical-ly reduces the rendering time for the same image quali-ty, making SurRender several orders of magnitude faster than general purpose image renderers, while still providing high physical accuracy. Various other opti-



mizations have also been implemented to reduce both noise (raytracing is a stochastic process) and rendering time.

Specific features for Entry, Descent and Land-

ing simulations: Planetary approach and landing mis-sions pose a specific challenge for an image simulation tool, which must handle planet-size texture at a very high resolution, in order to provide high-quality imag-es from the start of the approach until touch-down. SurRender is able to use “giant textures”, with a cus-tom protocol that maps the texture data to a virtual file up to 256 TB. Although the virtual file is accessed by the software like a regular file, with the same memory and cache optimizations, its content can be either dis-tributed on several files on a hard drive, or across the network. Alternatively, procedural texture generation can be used to generate texture data on-the-fly, thus allowing to generate image with a level of detail far beyond the original texture.

SuRender’s geometrical modelling of planetary bodies is based on an analytical ellipsoid shape, with custom flattening value, mapped with a DEM. In addi-tion to texture mapping described above, surface re-flectance properties can be finely modelled.

SurRender also provides depth maps and normal maps, that can be used as ground truth for algorithm validation, or used as input data for a Lidar model.

Fully customizable rendering chain: All ele-

ments of the simulation can be customized from sur-face materials (BRDF) to sensor models. SurRender comes with its own modelling tool, a high-level lan-guage derived from C, allowing to simply handle sca-lar and vectors types. This modelling language can be used to write models for materials, analytical shapes, or sensor electronics. Classical BRDF models are al-ready bundled with SurRender: lambertian model, Hapke model for moon and asterods, Oren-Nayar model for the jovian moons.

Figure 4: A: Rendering of Europa backlit by Jupi-ter, with increasing integration times. The amount of light per spectral band lit by Jupiter depends on the albedo and the colors of Jupiter’s clouds

Temporal and dynamic aspect of the simulation can

be handled, allowing to simulate physically effects

such as motion blur, or effects of a rolling-shutter de-tector on the acquired image.

Figure 5: HAS3 sensor simulation. Sub-windows at different integration times enable simultaneous observation of faint stars and bright objects.

Five years of use and constant improvements: SurRender’s development has started in 2011. Since then, it has been intensively used in many computer vision and space exploration projects by Airbus De-fence and Space and its partners. For instance, Sur-Render is currently used in R&D and technological development projects for the European Space Agency (JUICE, MREP, …), and for the European Community (NEOShield-2, H2020 program). It is also used to support the sizing and design of JUICE’s Navigation Camera.

It has also been used in Earth Observation imagery projects, where its has provided invaluable insight into the physics of image acquisition– including some un-expected second order effects – and enabled a deeper understanding of the impact of the acquisition parame-ters on the image quality.

New developments and features are foreseen for implementation within the next few months, among which the capability to simulate participating media (partially transparent objects) for finer atmostphere simulation. SurRender is also currently being coupled with an Optical Ground Support Equipment (OGSE) originally developed by Airbus Defence and Space to stimulate optically Star Trackers during ground testing. This coupling of SurRender with STOS (Star Tracker Optical Stimulator) is the first step of the development of a ground stimulation mean for Navigation Cameras, and will be tested in 2016 within a real-time test bench including a camera and a space-qualified CPU board.

References: [1] G.Jonniaux and D.Gherardi (2014), 9th ESA-GNC Conference



Documents

Poster Session Wednesday, June 15 from 5:00 to 7:00 PMippw2016.jhuapl.edu/docs/abstracts/Complete_Poster_session_abstr… · Poster Session Wednesday, June 15 from 5:00 to 7:00 PM